Extinguishable divert system

ABSTRACT

Various implementations of an extinguishable, solid propellant divert system for a flight vehicle are disclosed. Also disclosed are methods for using the divert system to control the flight of a flight vehicle. In one implementation, a divert system includes a hot gas generator pneumatically linked to one or more divert thrusters and an extinguishment valve. The extinguishment valve can be opened to rapidly depressurize the hot gas generator and extinguish the solid propellant burning inside. In another implementation, a method of controlling the trajectory of the flight vehicle includes repeatedly igniting and extinguishing the solid propellant in a hot gas generator and using the hot gas to provide divert thrust for the flight vehicle.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under one or more of thefollowing contracts awarded by the Missile Defense Agency through theDepartment of Defense (DoD) Small Business Innovative Research Program(SBIR). The U.S. government has certain rights in the invention.

Contract No. HQ0147-17-C-7424 (2017)

Contract No. HQ0147-16-C-7705 (2016)

Contract No. HQ0147-16-C-7706 (2016)

Contract No. HQ0147-15-C-7244 (2015)

Contract No. HQ0147-14-C-7873 (2013)

Contract No. HQ0147-13-C-7205 (2012)

Contract No. W91260-09-C-0008 (2009)

Contract No. W9113M-08-0069 (2008)

Contract No. W9113-07-C-0142 (2007)

Contract No. HQ0006-06-C-7479 (2006)

TECHNICAL FIELD

This relates to solid propellant divert control systems for flightvehicles that can be repeatedly extinguished and reignited to provideon-demand, multi-pulse, divert thrust over an extended operation time.It especially relates to such systems where the solid propellant isextinguished by rapid depressurization. The divert system can be part ofa solid propellant divert and attitude control system.

BACKGROUND

One of the greatest threats facing the world is the increasingproliferation of ballistic missiles and weapons of mass destruction.Despite reductions in the number of weapons deployed by the UnitedStates and the former Soviet Union, ballistic missile proliferationcontinues on a wide scale and could increase as technology istransferred to additional nations.

Nations invest in ballistic missiles because they provide the means toproject power both in a regional and strategic context and a capabilityto launch an attack from a distance. A nation without ballistic missilescan acquire them in a short period of time through an intensive researchand development program. In the future, ballistic missiles could evenbecome available to nonstate terrorist groups.

Missile defense technology being developed, tested and deployed by theUnited States is designed to counter ballistic missiles of allranges—short, medium, intermediate and long. Since ballistic missileshave different ranges, speeds, sizes, and performance characteristics,the ballistic missile defense system is an integrated, “layered”architecture that provides multiple opportunities to destroy missilesand their warheads before they reach their targets.

The missile defense system architecture includes: (1) networked sensors(including space-based) and ground and sea based radars for targetdetection and tracking; (2) ground and sea based interceptor missilesfor destroying a ballistic missile using either the force of a directcollision, called “hit-to-kill” technology, or an explosive blastfragmentation warhead; and (3) a command, control, battle management,and communications network providing operational commanders with theneeded links between the sensors and interceptor missiles.

One of the key components of the missile defense system is the standardmissile 3 (SM-3), the latest design of which is the SM-3 Block 1B. It isa ship and/or land-based missile used by the United States and itsallies to intercept short to intermediate range ballistic missiles aspart of the Aegis Ballistic Missile Defense System. Radar locates theballistic missile and the Aegis weapon system calculates a solution onthe target. Once a solution is in place, the missile is launched.

A solid fuel rocket booster launches the SM-3 out of a Mark 41 verticallaunching system (VLS). After launch, the missile establishescommunication with the launching platform (ship or ground installation)and proceeds towards the target. Once the booster or first stage burnsout, it detaches, and a second stage solid-fuel dual thrust rocket motor(DTRM) takes over propulsion through the atmosphere. The missilecontinues to receive mid-course guidance information from the launchingplatform and is aided by GPS data.

The second stage rocket motor eventually burns out and detaches and asolid-fuel third-stage rocket motor (TSRM) takes over propulsion. TheTSRM can propel the missile above the atmosphere if needed. The TSRM ispulse fired and provides propulsion for the SM-3 until approximately 30seconds to intercept when the TSRM separates from the kill vehicle (KV).

The KV includes a seeker system that searches for the target usingpointing data from the launching platform. The seeker system containssensors and/or other target acquisition components such as infrared (IR)sensors, radio frequency (RF) sensors, radar, and/or optics that areused to detect and pinpoint the location of the target. The seekersystem identifies the target and establishes a track for guidance.

A divert and attitude control system (DACS) such as the throttleabledivert and attitude control system (TDACS) used with the SM-3 Block 1Bmissile is used to maneuver the KV to the target. The DACS includes adivert system and an attitude control system (ACS). The divert systemremoves trajectory errors that remain after the earlier portions offlight and responds to guidance commands derived from seeker systemmeasurements. In general, the divert system provides the lateral motionfor the KV, and the ACS provides the angular control to stabilize KVpointing and control divert direction.

The DACS can maneuver the KV in various ways such as “diverting” thetrajectory of the KV or adjusting the attitude (pitch, roll, and yaw) ofthe KV. Divert movements are typically performed to move the KVlaterally or otherwise adjust the KV's trajectory. Attitude adjustmentsare performed to control the orientation of the KV with respect to aninertial frame of reference or another entity, which is usually thetarget. For example, the DACS can adjust the attitude of the KV toposition radar, optics, and other sensors towards the target. Divertmaneuvers typically require substantially more total impulse thanattitude adjustment maneuvers.

Although conventional DACS technologies, such as those used in the SM-3Block 1B TDACS, have functioned somewhat acceptably, they also sufferfrom a number of performance deficiencies in the following areas: (1)operating time (increased operating time is preferred), (2) energymanagement (on/off capability), (3) mass (less mass being preferred),and (4) divert distance. Accordingly, it would be desirable to provide aDACS system that improves one or more of these factors.

SUMMARY

A divert and attitude control system (DACS) includes an attitude controlsystem (ACS) and a divert system (DS). The DACS can be used with avariety of endoatmospheric and exoatmospheric flight vehicles. Forexample, it can be used as the DACS for the kill vehicle (KV) of amissile defense interceptor missile. It can also be used with any of theother flight stages of a guided missile. Also disclosed are methods ofusing the DACS to control the trajectory of a flight vehicle (divertsystem) and/or the attitude of the flight vehicle (attitude controlsystem).

The DACS can be implemented in various ways to realize one or more ofthe following potential advantages. One potential advantage is that theattitude control system and/or the divert system can provide on-demand,multiple thrust pulses and long duration operation. In someimplementations, for example, the attitude control system and/or thedivert system can provide thrust for at least 1000 seconds.

Another potential advantage of the DACS is that the attitude controlsystem and/or the divert system can use solid propellant. Solidpropellant is relatively stable and easy to transport compared to otheroptions such as hypergolic propellants. This makes it suitable for usewith flight vehicles launched from land or sea. Solid propellant is alsorelatively inexpensive compared to the alternatives. Another potentialadvantage is that the DACS is lightweight, which increases thecapabilities of the flight vehicle, particularly for a KV.

In some implementations, the attitude control system and/or divertsystem can use an extinguishable solid propellant. The propellant isignited to provide pressurized gas for the attitude thrusters and/ordivert thrusters. In some implementations, the attitude control systemand the divert system each include separate propellant. In oneimplementation, the propellant in the divert system is ignited by hotgas from the attitude control system.

In some implementations, the propellant in the attitude control systemand/or divert system is repeatedly extinguished and reignited. In oneimplementation, hot gas generated by burning the propellant in theattitude control system is used to repeatedly ignite the propellant inthe divert system. The solid propellant in the attitude control systemand/or the divert system can be extinguished in a number of ways. Insome implementations, the solid propellant is extinguished by rapiddepressurization.

In some implementations, the attitude control system can be usedindependently with divert systems other than those described in thisdocument. For example, the attitude control system can be used with athrottleable divert system such as the one currently used with the SM-3Block 1B missile.

In some implementations, the divert system can be used independentlywith attitude control systems other than those described in thisdocument. For example, the divert system can be used with an attitudecontrol system that uses high pressure, stored, cold gas to provideattitude thrust.

In some implementations, the DACS can provide continuous attitudecontrol and/or divert control capability for a relatively long period oftime. For example, the DACS can provide continuous attitude controland/or divert control capability for 100 seconds to 2000 seconds. Also,the DACS can provide continuous attitude control and/or divert controlcapability for at least 100 seconds, at least 200 seconds, at least 300seconds, at least 400 seconds, at least 500 seconds, and so forth.

In some implementations, the attitude control system and the divertsystem are separate components that are pneumatically linked. The twosystems are pneumatically linked in the sense that hot gas generatedfrom one system can be channeled to the other system to ignite thepropellant in the other system. In some implementations, hot gas fromthe attitude control system can be channeled to the divert system toignite the propellant in the divert system. In some implementations, hotgas from the attitude control system can be used to repeatedly ignitethe propellant in the divert system thereby eliminating the need forigniters in the divert system.

The attitude control system can have a variety of configurations. Insome implementations, the attitude control system includes a gasgenerator, an accumulator coupled to the gas generator, a valvepositioned between the gas generator and the accumulator, and attitudethrusters. The gas generator includes propellant that burns to providehot gas to the accumulator where it is stored. The accumulator ispneumatically linked to attitude thrusters that use the hot gas in theaccumulator to adjust the attitude of the flight vehicle.

The valve can be opened to recharge the accumulator with hot gas fromthe gas generator and, after the accumulator is full, closed to hold thepressurized hot gas in the accumulator. The valve can include variouscomponents that allow it to withstand the high temperatures and highpressures produced by the burning propellant. In one implementation, thevalve includes components made of a ceramic matrix composite such asC—ZrOC or C—SiC.

In some implementations, the valve extends at least part way into theaccumulator. In this configuration, the valve is pressurized when theaccumulator is recharged with hot gas. After the accumulator is full andthe valve is closed, the pressure inside the valve falls to ambientwhile the pressure in the accumulator remains. In this configuration,the pressure in the accumulator exerts hoop compression on the outsideof the valve.

In some implementations, the attitude control system includes a ventvalve that is in fluid communication with the gas generator and/or theaccumulator. The vent valve can be used to extinguish the propellant inthe gas generator. For example, after the accumulator is recharged bythe propellant burning in the gas generator, the valve to theaccumulator is closed and the vent valve is opened. The suddendepressurization in the gas generator extinguishes the propellant.

In some implementations, the attitude control system can operate in thefollowing manner. An initial propellant charge is positioned in theaccumulator and ignited with the valve closed. Hot gas fills theaccumulator until it reaches a set pressure or initial threshold atwhich the valve is opened. Hot gas flows from the accumulator to the gasgenerator and ignites the propellant in the gas generator for the firsttime. The gas generator produces additional hot gas and the pressuregradient reverses so that hot gas flows back into the accumulator.

The accumulator reaches a set point maximum pressure or upper thresholdpressure at which the valve to the accumulator is closed and the ventvalve is opened. The sudden depressurization in the gas generatorextinguishes the propellant. When the pressure in the accumulator dropsbelow a set point or lower threshold (due to attitude adjustments, etc.)or after a set amount of time, the accumulator is recharged by openingthe valve and closing the vent valve. Hot gas flows from the accumulatorto the gas generator and ignites the propellant. The hot gas from thegas generator pressurizes the accumulator and the cycle repeats itself.The accumulator can be recharged multiple times over the operationallife of the attitude control system.

In some implementations, the attitude control system is a low levelattitude control system (LLACS). For example, the attitude controlsystem that is part of the DACS for the KV can be a low level attitudecontrol system. The low level attitude control system can provideattitude control thrust throughout the final flight stage including whenthe divert system is active (burning propellant) or inactive(extinguished).

The divert system can have a variety of configurations. In someimplementations, the divert system includes a gas generator, an ignitionvalve, and divert thrusters. The ignition valve is positioned betweenthe divert system and the attitude control system to selectively allowhot gas from the attitude control system to enter the divert system andignite the propellant. For example, the ignition valve can be positionedbetween the accumulator and the divert gas generator. Opening theignition valve allows hot gas from the accumulator to contact and ignitethe propellant in the divert gas generator.

In general, the divert system typically includes substantially morepropellant than the attitude control system. This is because divertmaneuvers require substantially more thrust than attitude adjustments.In one implementation, the divert system includes at least 1.5× as muchpropellant as the attitude control system.

In some implementations, the divert system includes a first hot gasgenerator and a second hot gas generator spaced apart and positionedopposite each other along a lengthwise axis of the divert system. Thehot gas generators can be pneumatically linked to each other and to thedivert thrusters.

In some implementations, the one or more gas generators in the divertsystem are pneumatically separate from any propulsion rocket motors thatmay be included on the flight vehicle. This design makes it so thedivert system can produce hot gas when the propulsion rocket motors areoff or have separated from the flight vehicle.

A launch vehicle reaction control system (RCS) is also disclosed. TheRCS can be used to provide attitude control for small spacecraft launchsystems such as those that are used to launch satellites and the like.The RCS can be used in booster flyout attitude control systemapplications, post boost propulsion system (PBPS) applications, payloaddeltav applications, and attitude control systems to provide increasedsatellite orbital insertion accuracy.

The RCS can provide a number of advantages. One is that it can replaceconventional, heavier, and less performing cold gas attitude controlsystems resulting in a cost reduction up to 30% and a weight reductionup to 25%. It can provide a new affordable and higher performing solidRCS that fills an important technical gap for affordable access tospace.

The RCS can be used in a variety of applications including: hypersonicinflatable aerodynamic decelerator (HIAD), towed glider airLaunch system(TGALS), mars ascent vehicle reaction control system (MAV RCS),lunar/mars landers, large booster systems, low earth orbit smallsats,deep-space smallsats, non-propulsive gas generators, kinetic energy KVs,ground based strategic deterrent (GBSD) post boost propulsion andbooster roll control system applications, and/or hypersonic steeringapplications.

In some implementations, the RCS includes a gas generator, at least twothrusters, and one or more igniters. The gas generators hold solidpropellant that can be ignited with an igniter. The solid propellantburns generating hot gas. The thrusters are actuated to: (a) maintainthe pressure in the gas generate at a set level or within a certainrange and/or (b) exert thrust on the flight vehicle to adjust itstrajectory and/or attitude. The solid propellant can be extinguished byopening all or enough of the thrusters to cause rapid depressurizationof the gas generator. The solid propellant can be repeatedly ignited andextinguished to produce multiple thrust pulses.

It should be appreciated that referring to the gas generator as notbeing pneumatically linked to or pneumatically separate from apropulsion rocket motor means that no pneumatic pathway or linkageexists between the gas generator and the propulsion rocket motor. Ifsuch a pathway or linkage exists, then they are pneumatically linked andare not pneumatically separate. This is true regardless of the presenceor state (open/closed) of any valve that may be present in such apathway or linkage.

The systems, methods, and devices of this disclosure each have severalinnovative aspects, no single one of which is solely responsible for thedescribed desirable attributes. The summary is provided to introduce aselection of concepts in a simplified form that are further describedbelow in the detailed description. The summary and the background arenot intended to identify key concepts or essential aspects of thedisclosed subject matter, nor should they be used to constrict or limitthe scope of the claims. For example, the scope of the claims should notbe limited based on whether the recited subject matter includes any orall aspects noted in the summary and/or addresses any of the issuesnoted in the background.

DRAWINGS

The preferred and other implementations are disclosed in associationwith the accompanying drawings in which:

FIGS. 1-2 are conceptual diagrams of various implementations of a divertand attitude control system (DACS) including an attitude control systemand a divert system.

FIG. 3 is a perspective view of one implementation of the DACS in FIG.2.

FIG. 4 is a perspective view of the attitude control system from theDACS in FIG. 3.

FIGS. 5-8 are perspective views of a housing assembly from the DACS inFIG. 3. The housing assembly includes an accumulator valve, vent valve,divert valve, and a passage connecting all the valves.

FIG. 9 is a cross sectional view of the attitude control system showingthe inside of the accumulator and the accumulator valve.

FIG. 10 is a side view of the housing assembly in FIGS. 5-8 from theside of the accumulator valve.

FIG. 11 is a cross sectional view of the housing assembly in FIG. 10along line 10-10.

FIG. 12 is an aft of the housing assembly in FIGS. 5-8.

FIG. 13 is a cross sectional view of the housing assembly in FIG. 12along line 12-12.

FIG. 14 is a cross sectional perspective view of the housing assembly inFIG. 12 along perpendicular lines 13-13.

FIGS. 15A-15B are perspective views of one implementation of a throatarea of the accumulator valve.

FIG. 16 is a cross sectional perspective view of the vent valve.

FIG. 17 is a cross sectional perspective view of the divert valve.

FIG. 18 is a perspective view of another implementation of the DACS inFIG. 2.

FIG. 19 is a cross-sectional view of the divert ignition valve used withthe DACS in FIG. 18.

FIG. 20 is a partial cross-sectional perspective view of a prototypeattitude control system. The system includes an accumulator, gasgenerator, an accumulator valve positioned between the accumulator andthe gas generator, and a vent valve used to extinguish the gasgenerator.

FIG. 21 is a cross-sectional top view of the prototype attitude controlsystem with the major components delineated by dashed rectangles.

FIG. 22 is a cross-sectional view of the accumulator valve in theprototype attitude control system.

FIG. 23 is a partial cross-sectional perspective view of the accumulatorvalve in the prototype attitude control system.

FIG. 24 is a cross-sectional view of the vent valve in the prototypeattitude control system.

FIG. 25 is a cross-sectional view of the accumulator valve housingassembly in the prototype attitude control system.

FIG. 26 is a cross-sectional view of the accumulator housing in theprototype attitude control system.

FIG. 27 is a cross-sectional view of the gas generator in the prototypeattitude control system.

FIG. 28 is a graph of the test data produced by a first hot fire of theprototype attitude control system. The graph shows the pressure in theaccumulator and gas generator as well as the actuation of theaccumulator valve and the vent valve.

FIG. 29 is a detailed graph of the data in FIG. 28 for the first 7.25seconds of the hot fire test, which includes the initial pressurizationand first recharge of the accumulator.

FIG. 30 is a graph of the test data produced by a second hot fire of theprototype attitude control system. The graph shows the pressure in theaccumulator and gas generator as well as the actuation of theaccumulator valve and the vent valve.

FIG. 31 is a graph of the test data produced by a third hot fire of theprototype attitude control system. The graph shows the pressure in theaccumulator and gas generator as well as the actuation of theaccumulator valve and the vent valve.

FIGS. 32 and 33 are perspective views of a prototype divert system fromthe forward end and aft end, respectively.

FIG. 34 is a perspective view of a partial cross section of theprototype divert system in FIGS. 32 and 33.

FIG. 35 is a cross sectional view along a lengthwise axis of theprototype divert system in FIGS. 32 and 33.

FIG. 36 is a cross sectional view of the central area between the gasgenerators of the prototype divert system in FIGS. 32 and 33.

FIG. 37 is a perspective view of the gas generator from the prototypedivert system in FIGS. 32 and 33.

FIG. 38 is a cross sectional view of the gas generator from theprototype divert system in FIGS. 32 and 33.

FIG. 39 is a perspective view of a cross section of the gas generatorfrom the prototype divert system in FIGS. 32 and 33.

FIG. 40 is a cross sectional view of the interface between the case andthe dome from the gas generator of the prototype divert system in FIGS.32 and 33.

FIG. 41 is a cross sectional view of a hot gas tube connecting andextending between the gas generators of the prototype divert system inFIGS. 32 and 33.

FIG. 42 is a cross sectional view of an igniter coupled to the gasgenerator of the prototype divert system in FIGS. 32 and 33.

FIG. 43 is a perspective view of a divert thruster in the prototypedivert system in FIGS. 32 and 33.

FIG. 44 is a perspective view of a cross section of the divert thrusterin FIG. 43.

FIG. 45 is a cross sectional view of the divert thruster in FIG. 43coupled to the gas generator.

FIG. 46 is a cross sectional view of attitude thrusters coupled to theprototype divert system in FIGS. 32 and 33.

FIG. 47 shows the results of a thermal analysis of the prototype divertsystem. It includes a chart of the temperature at various locationsalong with an illustration of the locations.

FIG. 48 is a perspective view of a reaction control system for a launchvehicle.

FIG. 49 is a perspective view of the reaction control system in FIG. 48with the outer housing removed.

FIG. 50 is a cross sectional view of the reaction control system in FIG.48 coupled to the after of a nose cone of a launch vehicle.

FIG. 51 is a perspective view of a partial cross section of the thrustsystem from the reaction control system in FIG. 48.

FIG. 52 is an aft or bottom perspective view of the thrust system fromthe reaction control system in FIG. 48.

FIG. 53 is a partially exploded top perspective view of the thrustsystem in FIG. 52.

FIG. 54 is a perspective view of an igniter that can be used with thereaction control system in FIG. 48.

FIG. 55 is a cross-sectional view of the igniter in FIG. 54.

FIG. 56 is a perspective view of a cross section of the igniter in FIG.54.

FIG. 57 is a cross sectional view of a thruster from the reactioncontrol system in FIG. 48.

FIG. 58 is a perspective view of another reaction control system for alaunch vehicle.

FIG. 59 is a perspective view of the reaction control system in FIG. 58with the outer housing removed.

FIGS. 60-63 illustrate the operation of a reaction control system.

DETAILED DESCRIPTION

Overview

FIGS. 1-2 show conceptual diagrams of various implementations of adivert and attitude control system (DACS) 10. The DACS 10 can be used ina variety of ways and with a variety of flight systems or vehicles 11.In some implementations, the DACS is included as part of a missiledefense interceptor missile launched to destroy a target such as aballistic missile. For example, the DACS 10 can be used during the finalstage of flight to maneuver a kill vehicle (KV) into the target. TheDACS 10 can also be used for advanced upper stage booster divert and/orattitude control applications.

In one implementation, the DACS 10 can be included as part of thestandard missile 3 (SM-3) used in current missile defense systems. Forexample, the DACS 10 can be part of the final stage control system thatmaneuvers the KV into the target. The DACS 10 can also be used with anyof the other stages of the SM-3. For example, the DACS 10 can be usedwith the third stage rocket motor of the SM-3 to perform divert andattitude adjustment maneuvers.

The DACS 10 can be used independently or in conjunction with apropulsion rocket motor. For example, in some implementations, the KVdoes not include a propulsion rocket motor. In such situations, the DACS10 can be used exclusively to maneuver the KV during the final stage offlight—e.g., adjust the attitude and trajectory of the KV to interceptthe target. In some implementations, the DACS 10 is pneumaticallyseparate from any propulsion rocket motors that may be included on theflight vehicle. This design makes it so the divert system can producehot gas when the propulsion rocket motors are off or have separated fromthe flight vehicle.

In some implementations, the DACS 10 uses hot combustion gas to providethrust for both divert and attitude adjustment maneuvers. This isespecially advantageous in the context of attitude adjustments. Thistype of system can provide a greater amount of thrust than systems thatuse pressurized cold gas for attitude adjustments, which gas must beprovided as a pre-pressurized container that is launched with the flightvehicle 11. Also, a hot gas system is safer to store, transport, andhandle than high pressure containers.

In some implementations, the DACS 10 generates and stores the hot gas.The pressures produced by this process can be significant. In oneimplementation, the DACS 10 can withstand a maximum pressure of at least1,000 psia, at least 1,500 psia, at least 2,000 psia, at least 2,500psia, at least 3,000 psia, or at least 3,500 psia. In anotherimplementation, the DACS 10 is designed to withstand a maximum pressureof 1,000 to 3,500 psia, 1,500 psia to 3,000 psia, or 2,000 psia to 3,000psia.

In some implementations, the DACS 10 is a solid propellant DACS (SDACS).This means that the DACS 10 burns solid propellant to provide thrust fordivert and attitude adjustment maneuvers. In general, it is preferableto use solid propellant because it is safer to store, handle, andtransport than liquid propellant.

In some implementations, the solid propellant can be extinguishable.This makes it possible to repeatedly ignite and extinguish thepropellant during operation, which increases the operational time of theDACS 10. In some implementations, the solid propellant can beextinguished by sudden rapid depressurization. In general, rapiddepressurization involves reducing the pressure to ambient within nomore than 1 second, no more than 750 ms, no more than 500 ms, or no morethan 250 ms.

The DACS 10 can be extinguished and reignited any suitable number oftimes. In some implementations, the DACS 10 can be extinguished andreignited (or ignited and extinguished) at least 20 times duringoperation, at least 25 times during operation, or at least 30 timesduring operation.

The DACS 10 can operate for a relatively long period of time. Theoperational time of the DACS 10 is the period after it is initiallyignited during which it can supply thrust for divert and attitudeadjustment maneuvers. The operational time can include periods when theDACS 10 is extinguished. In general, it is desirable to maximize theoperational time of the DACS 10 given the constraints of the particularflight vehicle. Long duration operation allows the flight vehicle totravel longer distances and operate with greater efficiency.

In one implementation, the DACS 10 has an operational time of at least100 seconds, at least 200 seconds, at least 300 seconds, at least 400seconds, at least 500 seconds, at least 600 seconds, at least 700seconds, at least 800 seconds, at least 900 seconds, or at least 1000seconds. In another implementation, the DACS 10 has an operational timeof 100 to 2,000 seconds.

In some implementation, the DACS 10 can use solid propellant andsatisfies one or more of the specifications shown in Table 1.

TABLE 1 Solid DACS Specifications Parameter Value Operating time ≥300 sOperating mode Extinguishable and/or throttling Ignition criteria Hotgas storage is ≥500 psia within 0.5 s of ignition Maximum mass Themaximum mass, when fully loaded with expendables, shall not exceed 30.0lb. AIAA Standard 120-2006 shall be used as a guideline for applyingmass growth and margin. Maximum At least 15 DACSs shall fit inside anSM-3 BLKIIA envelope shroud. Thruster ACS usage shall account forunbalanced divert and misalignment ACS thrust and all thrustermisalignment, both by design and manufacturing error. Structuralstiffness The KV first body bending will be greater than 300 Hz in freeflight, including payload interfaces. Assume a rigid payload. Payloadinterface Maximum payload interface temperature will not temp exceed250° F. during operation. Maximum payload Payload maximum mass shall notexceed 9.0 lb. mass Split payload mass Split payload mass allocationshall be 56.77% allocation (forward) and 43.23% (aft). Split payload Thesplit payload CG's are located 1.97 in forward center-of-gravity of theforward interface and 1.89 in aft of the aft interface. Split payloadThe split payload moments of inertia about the moments of inertiapayload's CG's shall be as follows (all units in lbm · in²): Front: Ixx= 148.8, Iyy/Izz = 188.5 Aft: Ixx = 148.8, Iyy/Izz = 79.4 Normal tempSystem shall meet performance and reliability after continuous exposureto assembly and check out temperature range of +50° F. to +95° F. duringthe final assembly and missile encanisterization period of up to 180days. Air conditioning System shall meet performance and reliabilityafter malfunction temp continuous exposure to assembly and check outtemperature range of 0° F. to +120° F. due to air conditioningmalfunction for a period not to exceed 72 hours. There can be up tothree air conditioning malfunctions during any 180 day period. Storagetemp System shall meet performance and reliability after continuousexposure to storage temperature range of −20° F. to +120° F. for periodsup to 2 years. The conditions apply with maximum variation in any 24 hrperiod of 30° F. Transportation System shall meet performance andreliability after temp exposure to transportation temp range of −20° F.to +130° F. for periods of up to 5 days. Normal Pressure System shallmeet performance and reliability after exposure to assembly and checkoutatmospheric pressures of 15.4-11.3 psia (sea-level to 7000 ft). PHS&TPressure System shall meet performance and reliability after exposure toPHS&T atmospheric pressures of 15.4- 2.7 psia (sea-level to 40,000 ft).

Referring back to FIG. 1, the DACS 10 includes an attitude controlsystem 12 (alternatively referred to as an attitude control subsystem)and a divert system 14 (alternatively referred to as a divertsubsystem). It should be appreciated that the description of the DACS 10above applies equally to one and/or both systems 12, 14. For example,the description of the operational time of the DACS 10 is applicableindividually to the attitude control system 12 and/or the divert system14.

The attitude control system 12 and the divert system 14 described inthis document do not need to be used together although doing so is oftenpreferable. The attitude control system 12 can be used with other divertsystems. For example, the attitude control system 12 can be used with anoff-the-shelf divert system 14 that is adapted to work with the attitudecontrol system 12.

Likewise, the divert system 14 can be used with other attitude controlsystems. For example, the divert system 14 can be used with anoff-the-shelf attitude control system I that is adapted to work with thedivert system 14. In some implementations, the divert system 14 can beused with an attitude control system that uses cold gas to providethrust.

The attitude control system 12 includes an accumulator 16, a gasgenerator 18, an accumulator valve or first valve 20, a vent valve,extinguishment valve, or second valve 22, and one or more attitudethrusters 24. The divert system 14 includes a divert valve or divertignition valve 26, a gas generator 28, and divert thrusters 30.

In some implementations, the systems 12, 14 are physically separateunits coupled together to form the DACS 10 as shown in FIGS. 3 and 18.For example, each system 12, 14 can include its own propellant (notshown), thrusters 24, 30, and the like. In some implementations, thesystems 12, 14 are pneumatically linked so that hot gas from theattitude control system 12 can be used to ignite the propellant in thedivert system 14 one or more times. The divert valve 26 can be used tocontrol the flow of hot gas between the attitude control system 12 tothe divert system 14. In some other implementations, the systems 12, 14are not pneumatically linked and operate separately.

It should be appreciated that the boundaries between the systems 12, 14as depicted in the FIG. 2 are conceptual in nature and subject to changedepending on the circumstances. For example, the divert valve 26 isshown as part of the divert system 14 in FIG. 2. However, the divertvalve 26 could also be considered part of the attitude control system 12if it is produced as part of the same unit that includes the componentsof the attitude control system 12. Alternatively, the divert valve 26could be part of the unit that includes the components of the divertsystem 14. Likewise, the vent valve 22 is shown as part of the attitudecontrol system 12 when it could just as easily be considered part of thedivert system 14.

It should be appreciated that divert maneuvers require more force thanattitude adjustments. Accordingly, the divert system 14 is generallylarger than the attitude control system 12. In one implementation, thedivert system 14 includes substantially more propellant than theattitude control system 12. For example, the divert system 14 caninclude 1.5× to 10× as much propellant, or more, as the attitude controlsystem 12. The divert system 14 can also provide more total impulse thanthe attitude control system 12. For example, the divert system 14 canprovide 1.5× to 10× as much total impulse, or more, than the attitudecontrol system 12.

One implementation of the DACS 10 is shown in FIG. 3. The accumulator 16has a circular or toroidal shape that encircles the base of the divertsystem 14. The attitude control system 12 includes a pair of housingassemblies 32 coupled to opposite sides of the accumulator 16. Thehousing assemblies 32 extend upward from the accumulator 16 adjacent tothe outside of the divert system 14. The upper end of each housingassembly 32 is coupled to a gas generator 18. The attitude thrusters 24are pneumatically linked to the accumulator 16 and extend upward nearthe forward end of the DACS 10.

Another implementation of the DACS 10 is shown in FIG. 18. It is roughlysimilar to that shown in FIG. 18. One notable change is the position ofthe attitude thrusters 24. They are positioned closer to and aft of theaccumulator 16. Another notable change is that the DACS 10 is that thevent valves 22 and the divert ignition valve 26 have been relocated frombeing underneath the gas generators 18 to being on top of the gasgenerators 18. The divert system 14 also includes two gas generators 28spaced apart along a longitudinal axis 36 of the divert system.

The details of the attitude control systems 12 and the divert systems 14shown in FIGS. 3 and 18 are described in greater detail below.

Attitude Control System (ACS)

FIG. 4 shows the attitude control system 12 in FIG. 3 separately fromthe divert system 14. Each housing assembly 32 includes an accumulatorvalve 20, a vent valve 22, a divert valve 26, and one or more passages34 pneumatically linking the gas generator 18 and the valves 20, 22, 26.The passages 34 (FIGS. 9, 11, 13, and 16-14) allow hot gas to flow fromthe gas generator 18 to the valves 20, 22, 26. In this manner, the gasgenerator 18 and the valves 20, 22, 26 are pneumatically linked to eachother. Perspective views of the housing assembly 32 are shown in FIGS.5-8.

The accumulator valve 20 controls the flow of hot gas between the gasgenerator 18 and the accumulator 16. The vent valve 22 is used to causea rapid depressurization of the gas generator 18 to extinguish thepropellant burning inside. The divert valve 26 is used to selectivelyallow hot gas to flow into the divert system 14 and ignite thepropellant for divert maneuvers. The valves 20, 22, 26 are operated withactuators 38, 40, 42, respectively.

In general, it is desirable to provide a single accumulator 16 eventhough the attitude control system 12 can include more than one of theother components. The reason a single accumulator 16 is advantageous isbecause it equalizes the pressure of the hot gas supplied to theattitude thrusters 24. If two accumulators 16 were used, then itincreases the likelihood of a pressure differential between theaccumulators 16, which could increase the variability of the thrustprovided to individual attitude thrusters 24.

Despite the advantages of a single accumulator 16, it should beappreciated that other implementations can include multiple accumulators16. For example, multiple accumulators 16 can be used if eachaccumulator is coupled to an independent set of thrusters that aren'tdesigned to function together in a concerted manner.

In some implementations, the attitude control system 12 is positionedsymmetrically along a lengthwise axis 36 of the flight vehicle 11 or ofthe DACS 10. In the implementation shown in FIG. 3, the lengthwise axis36 extends through the center of the toroidal shape of the accumulator16 and through the center of the divert system 14. A symmetrical designis advantageous because it evenly distributes the weight of the attitudecontrol system 12, which helps stabilize the flight vehicle duringflight.

In some implementations, the weight of the attitude control system 12remains symmetrical throughout operation. The weight of the attitudecontrol system 12 changes as propellant is burned in the gas generators18. In the implementation shown in FIG. 3, the propellant is distributedequally in the gas generators 18 so that as it burns, the center ofgravity of the attitude control system 12 shifts forward along thelengthwise axis 36 but doesn't shift side to side.

It should be appreciated that the attitude control system 12 can haveany suitable shape and/or configuration. For example, the accumulator 16can have a cylindrical, hexagonal, or other shape. Also, the attitudecontrol system 12 can include a single housing assembly 32 with a singlegas generator 18, accumulator valve 20, vent valve 22, and divert valve26. In other implementations, the attitude control system 12 can includethree or more housing assemblies 32 with a corresponding number of gasgenerators 18 and valves 20, 22, 26.

In some implementations, the attitude control system 12 can withstandthe same pressures and operate for the same amount of time as the DACS10. In general, it should be appreciated that any individual parameterdisclosed in connection with the DACS 10 also applies to the attitudecontrol system 12. For example, if the DACS 10 can withstand a givenpressure or temperature, then the attitude control system 12 canwithstand the same pressure or temperature. Also, the operational timesof the DACS 10 apply equally to the attitude control system 12.

In some implementations, the attitude control system 12 is a stand-aloneunit that can be used with any suitable divert system 14. The divertvalve 26 can be considered part of the attitude control system 12 inthese implementations or the divert valve 26 can be eliminated. In thosesituations where the divert valve 26 is included, it can be coupled tothe divert system 14 to pneumatically link the two systems 12, 14. Thestand-alone nature of the attitude control system 12 makes it flexibleand easy to adapt to various divert systems 14 and flight vehicles.

The attitude control system 12 can operate in a variety of differentways. In some implementations, the attitude control system 12 operatesas follows. An initial charge of propellant or, in other words, a startgrain of propellant is positioned in the accumulator 16. The accumulatorvalve 20 is closed to isolate the accumulator 16 from the othercomponents in the attitude control system.

The initial charge is ignited to activate the attitude control system 12and pressurize the accumulator 16. The amount of propellant in theinitial charge is sufficient to pressurize the accumulator 16 above aninitial set point or initial threshold. The initial set point can be anysuitable minimal pressure level. In one implementation, the initialcharge pressurizes the accumulator 16 to at least 300 psia, at least 400psia, at least 500 psia, or at least 600 psia.

Once the pressure in the accumulator 16 reaches the initial set point,the accumulator valve 20 is opened to allow the hot gas to flow throughthe passages 34 in the housing assembly 32 to the gas generator 18. Thehot gas ignites the propellant in the gas generator 18, which causes thepressure to continue to rise in the housing assembly 32 and theaccumulator 16 until it reaches an upper threshold or first set point.It should be noted that the vent valve 22 and the divert valve 26 areclosed up to this point.

The maximum pressure can be set at any suitable amount. In oneimplementation, the maximum pressure is no more than 4,000 psia, no morethan 3,500 psia, no more than 3,000 psia, no more than 2,500 psia, or nomore than 2,000 psia. When the pressure in the accumulator 16 reachesthe upper threshold, the accumulator valve 20 is closed to keep thepressurized hot gas in the accumulator 16. At the same time, the ventvalve 22 is opened to rapidly depressurize the gas generator 18 andextinguish the propellant. The vent valve 22 remains open until theaccumulator 16 is recharged to ensure that the propellant is fullyextinguished.

In order to extinguish the propellant, the pressure in the gas generator18 should drop rapidly to ambient levels (in space this is 0 psia). Insome implementations, the pressure drops to ambient levels within nomore than 1 second, no more than 750 ms, no more than 500 ms, or no morethan 250 ms.

The accumulator 16 is now in a fully charged or fully pressurizedcondition. The hot gas in the accumulator 16 is released through theattitude thrusters 24 as attitude adjustments are made to the flightvehicle. The accumulator 16 is recharged when a lower threshold orsecond set point is reached. The second set point can be a minimumpressure in the accumulator 16, a set amount of time since the lastrecharge, or both. In one implementation, the accumulator 16 isrecharged when either the pressure falls below a lower threshold or aset amount of time has passed since the last recharge.

In some implementations, the accumulator 16 is recharged when thepressure drops below 1,000 psia, below 750 psia, or below 500 psia. Inother implementations, the accumulator is recharged after 2 seconds, 3seconds, 5 seconds, 10 seconds, 20 seconds, 30 seconds, 40 seconds, or45 seconds.

In some implementations, the accumulator 16 can be recharged more oftenat the beginning of the process to heat up the system hardware tooperating temperature. In other words, the set amount of time betweenrecharges can be lower initially and then increased as the system 12heats up. The hardware absorbs heat from the hot gas. If it absorbs toomuch heat, then the hot gas may not successfully ignite the propellantin the gas generator 18.

The accumulator 16 is recharged by closing the vent valve 22 and openingthe accumulator valve 20. Hot gas from the accumulator 16 flows to thegas generator 18 and ignites the propellant. The process of pressurizingthe accumulator 16 described above is repeated.

It should be appreciated that the accumulator 16 can be recharged manytimes during the operational life of the attitude control system 12. Insome implementations, the accumulator 16 is recharged at least 20 times,or at least 25 times. Repeatedly igniting and extinguishing thepropellant in the gas generator 18 helps to extend the operation time ofthe attitude control system 12.

The configuration of the attitude control system 12 provides a number ofadvantages. One advantage is that the attitude control system 12 onlyneeds a single igniter for its entire operational life. Once theaccumulator 16 is initially pressurized, the hot gas contained in it canbe used for all subsequent propellant ignitions in either or both of theattitude control system 12 and the divert system 14. This is in contrastto conventional solid propellant systems, which require a separateigniter each time the propellant is reignited.

Another advantage is that the attitude control system 12 complies withMIL-STD-1901A, which is the safety criteria for the design of munitionrocket and missile motor ignition systems. One of the reasons the designof the attitude control system 12 is compliant is because the igniterand initial charge of propellant are separated from the propellant inthe gas generator 18 and the propellant in the divert system 14. Thismeans that during storage and handling the attitude control system 12can be configured so that if the initial charge accidentally ignites itwon't ignite the other propellant.

In one implementation, the attitude control system 12 can be stored withthe accumulator valve 20 and the vent valve 22 open. In this state, thehot gas produced by an accidental ignition of the initial charge isimmediately vented through the vent valve 22. The hot gas cannot produceenough pressure to ignite the propellant in either the gas generator 18or the divert system 14.

In another implementation, the attitude control system 12 can be storedwith the accumulator valve 20 closed and the attitude thrusters 24 open.In this implementation, the hot gas produced by an accidental ignitionof the initial charge is immediately vented through the attitudethrusters 24. In yet another implementation, the attitude control system12 can be stored with the accumulator valve 20, the vent valve 22, andthe attitude thrusters 24 open. Numerous other configurations are alsopossible.

Low Level Attitude Control System (LLACS)

In one implementation, the attitude control system 12 is a low levelattitude control system designed specifically for use with the SM-3interceptor missile. For example, the attitude control system 12 can beused to adjust the attitude of the kill vehicle during the final stageof flight just before it impacts the target.

In one implementation, the KV includes a seeker system having varioussensors, transmitters, and/or receivers that allow it to send andreceive information. For example, the sensors can be used to obtaininformation about the target from heat signatures, light emissions,radio wave emissions, and the like. In some implementations, the sensorscan be used to find and track the heat signature of the target. Theattitude control system 12 can be used to adjust the attitude of the KVto point the sensor directly at the target. The attitude control system12 can be used in numerous other ways as well.

In some implementations, the attitude control system 12 can be a smallcompact system that is limited in the amount of total impulse it canprovide. For example, it can be configured to provide no more than 800lbf·s of impulse, no more than 600 lbf·s of impulse, no more than 400lbf·s of total impulse, or no more than 300 lbf·s of total impulse.

In some implementations, the low level attitude control system satisfiesone or more of the specifications in Table 2 and/or Table 3. A low levelattitude control system meeting these requirements may be especiallysuitable for use with the SM-3's KV.

TABLE 2 Low Level Attitude Control System Specifications 1 ParameterValue Min. pressure   500 psia Nominal max. pressure 3,000 psia Rechargecycles ≥28 Max. expected operating 3,500 psia pressure (MEOP) Structuralfactors of safety FS_(ULT) = 1.25; FS_(YLD) = 1.10; FX_(PRF) = 1.0 atMEOP Configuration/layout Common accumulator; dual gas generators andhousing assemblies Delivered total low level ≥200 lbf · s (≥100 lbf · sper accumulator impulse valve) Thruster(s) inlet temperature ≤2000° F.SDACS ignition capability Pressurize 200 in³ volume to ≥500 psia in ≤0.5s System weight ≤10 lbm Propellant type Extinguishable Ignition systemsafety MIL-STD-1901A compliant

TABLE 3 Low Level Attitude Control System Specifications 2 ParameterValue Operation time Launch to Initiation: ≤240 s Upon initiationcommand: ≥1000 s ACS minimum ACS shall be capable of rotating the KV 2.0angular velocity degrees and provide stabilization within 0.5 s inpitch, yaw, and roll directions ACS pointing ACS shall be capable ofstabilizing the KV accuracy within 0.02 degrees of the commanded valueACS coning motion Unknown; see total ACS impulse requirement as point ofdeparture ACS coning angular Unknown; see total ACS impulse requirementvelocity as point of departure Total ACS impulse ACS impulse shall be≥80 lbf · s ACS minimum <0.045 lbf · s during divert pulses impulse bit<0.00044 lbf · s during low-level operation ACS control authority ACScontrol authority shall be no less than 1.5 for the entire operation ACSduty cycle ACS shall fire as needed during operation time ACS pulseoperation Each pulse shall range from 0.005 s to time continuousResponse time (from Response time from 0 to 90% thrust command ignitioncommand) from an ignition command shall be ≤2.0 s Response time (whenResponse time from 0 to 90% thrust command already ignited) from athrust command shall be: Less than 4 ms during divert pulses Less than 3ms during low level operation

Each of the components of the attitude control system 12 are describedin greater detail as follows. The components can be off-the-shelf partsor custom manufactured for a specific application. The components thatare subject to the most extreme conditions are more likely to be custommanufactured.

Accumulator

The accumulator 16 can have any suitable configuration. In general, theaccumulator 16 is in the form of an enclosure that is capable of holdingthe hot gas generated by burning the solid propellant. The accumulator16 can have a variety of shapes including those described above. Theaccumulator 16 can also have any number and variety of interface ports.

The accumulator 16 can have any suitable amount of internal free volume.A larger amount of free volume means that the accumulator 16 does notneed to be recharged as often. However, it also means that theaccumulator 16 weighs more. Thus, there is a trade-off between internalfree volume and weight. In one implementation, the accumulator 16includes at least 20 in³ of internal free volume, at least 25 in³ ofinternal free volume, at least 30 in³ of internal free volume, at least35 in³ of internal free volume, at least 40 in³ of internal free volume,at least 45 in³ of internal free volume, or at least 50 in³ of internalfree volume.

The accumulator 16 can be made of any suitable material that is capableof withstanding the high temperatures and high pressures produced by thehot gas. In some implementations, the accumulator 16 is made ofstainless steel or a stainless steel alloy. For example, the accumulator16 can be made of 17-4 H1150 stainless steel alloy. In otherimplementations, the accumulator 16 can be made of titanium.

In some implementations, the accumulator 16 satisfies one or more of thespecifications set forth below in Table 4. This design of theaccumulator 16 may be especially suitable for use with a low levelattitude control system.

TABLE 4 Accumulator Specifications Parameter Value Internal free volume≥50 in3 Configuration Toroidal Interface ports 2x valve ports; 2xigniters; 2x thruster outlets, 1x pressure transducer Operatingpressure/MEOP 500 to 3,000 psia/3,500 psia Factors of safety at MEOPFS_(ULT) = 1.25; FS_(YLD) = 1.10; FX_(PRF) = 1.0

Gas Generator

The gas generator 18 is coupled to a forward or first end 44 of thehousing assembly 32. In general, the gas generator 18 is a containerconfigured to hold the propellant during storage and operation of theattitude control system 12. It should be appreciated that the gasgenerator 18 can have any suitable size and shape.

In some implementations, the gas generator 18 is a cylindrical canister.One end of the canister is coupled to the forward end 44 of the housingassembly 32. In other implementations, the gas generator 18 can have aspherical, hexagonal, or other shape. The gas generator 18 can be madeof any suitable material. In general, the gas generator 18 should becapable of withstanding the temperatures and pressures associated withcombustion of the propellant. In some implementations, the gas generator18 can be made of the same material as the accumulator 16.

The gas generator 18 can include any type of propellant. In oneimplementation, the propellant is solid propellant. In anotherimplementation, the propellant is extinguishable. In yet anotherimplementation, the propellant is an extinguishable, solid propellant.The propellant can be purchased commercially as an off-the-shelf productor custom designed for use with the gas generator 18.

In some implementations, the gas generator 18 satisfies one or more ofthe specifications set forth below in Table 5. This design of the gasgenerator 18 may be especially suitable for use with a low levelattitude control system.

TABLE 5 Gas Generator Specifications Parameter Value Max. propellantgrain 2.6 inches diameter Internal free volume ≥2 in³ (includesplumbing) Propellant type Extinguishable Operating pressure/ 500 to3,000 psia/3,500 psia MEOP Factors of safety at FS_(ULT) = 1.25;FS_(YLD) = 1.10; FX_(PRF) = 1.0 MEOP

Accumulator Valve

The accumulator valve 20 moves between an open position where hot gascan flow into and out of the accumulator 16 and a closed position wherehot gas is prevented from flowing into and out of the accumulator 16.The accumulator valve 20 is shown in the open position in FIGS. 9,11,and 13-14.

The accumulator valve 20 is subject to some of the harshest conditionsin the attitude control system 12. It is one of the few components thatis subjected to high temperatures and high pressures for the entireduration of the operation of the attitude control system 12. Most of theother components have an opportunity to cool off at one point oranother. The high temperatures and high pressures place a tremendousamount of stress and strain on the accumulator valve 20.

It should be appreciated that in some implementations, the accumulatorvalve 20 can be an off-the-shelf valve or can be adapted from anoff-the-shelf valve. For example, an off-the-shelf valve may be suitablefor situations having relatively lower temperatures and pressures andwhen the attitude control system 12 isn't a mission critical component.In other implementations, the accumulator valve 20 can be customdesigned for the specific application.

The accumulator valve 20 seals the accumulator 16 shut between rechargecycles. The accumulator valve 20 should not leak more than a minor orinsubstantial amount. If the accumulator valve 20 leaks more than this,then the accumulator 16 will need to be recharged more often and the gasgenerator 18 will need to be enlarged to hold more propellant, both ofwhich are undesirable.

FIGS. 9-14 show various cross-sectional views of the accumulator valve20. The accumulator valve 20 also includes a poppet 50, a poppet guide52, a valve shaft 54, and a valve shaft adapter 56. These componentsmove lengthwise (axially) inside the accumulator valve 20 to open andclose it.

The accumulator valve 20 includes a first or proximal end 58 and asecond or distal end 60. The accumulator valve 20 includes an actuatorseal plate 70 positioned at the second end 60. The actuator 38 iscoupled to the actuator seal plate 70. The actuator seal plate 70prevents the hot gas from escaping through the second end 60 of theaccumulator valve 20.

The actuator 38 engages the valve shaft adapter 56 at the second end 60of the accumulator valve 20. The actuator 38 opens the accumulator valve20 by pushing the valve shaft 54 lengthwise towards the first end 58.The valve shaft 54 contacts and pushes the poppet guide 52 lengthwise,which, in turn, pushes the poppet 50 open. In one implementation, thepoppet 50 is coupled to and moves in tandem with the poppet guide 52.

In some implementations, the only way to close the accumulator valve 20is with the force of the pressure in the accumulator 16. The actuator 38only opens the accumulator valve 20; it doesn't close it. After theinitial charge has pressurized the accumulator 16, the actuator 38 opensthe accumulator valve 20 to allow hot gas to flow to the gas generator18. In this state, the pressure is higher in the accumulator 16 andlower in the gas generator 18 creating a pressure gradient from theformer to the latter. The actuator 38 holds the accumulator valve 20open as the hot gas flows from the accumulator 16 to the gas generator18.

When the propellant ignites, the pressure gradient reverses so that thepressure is higher in the gas generator 18 and lower in the accumulator16 causing the hot gas to flow in the opposite direction. The actuator38 no longer holds the accumulator valve 20 open. Instead, the flow ofhot gas holds it open. When the accumulator 16 is fully recharged, thevent valve 22 opens causing the pressure gradient to reverse again. Hotgas flows from accumulator 16 to the vent valve 22. The actuator 38moves the valve shaft 54 lengthwise back towards the second end 60 ofthe accumulator valve 20 and the flow of hot gas pushes the poppet 50closed.

In one implementation, the valve shaft 54 only contacts the poppet guide52 when the accumulator valve 20 is open. When it is closed, the valveshaft 54 is retracted towards the second end 60 of the accumulator valve20 far enough that it no longer contacts the poppet guide 52. Thisprovides a thermal break between the valve shaft and the poppet guide52, which reduces the heat load on the actuator 38 thereby extending itsuseful life.

It should be appreciated that the poppet 50, poppet guide 52, valveshaft 54, and valve shaft adapter 56 can be made of any suitablematerial. All of these components are subjected to high temperatures,especially the first three, and should be made of materials that arecapable of withstanding the temperatures. In some implementations, thepoppet 50 can be made of rhenium molybdenum and the poppet guide 52 andthe valve shaft 54 can be made of a ceramic matrix composite.

In some implementations, the accumulator valve 20 includes a shield orshaft shield 74 that surrounds the valve shaft 54. The shield 74 can bemade of any suitable high temperature resistant material such as rheniummolybdenum.

The accumulator valve 20 includes a main body 48 through which the hotgas flows. The main body 48 is positioned in a valve housing 62. A layerof main body insulation 64 is provided between the valve housing 62 andmain body 48 near the first end 58 of the accumulator valve 20, which isthe area that gets the hottest. The main body insulation 64 preventsheat transfer from the main body 48 to the valve housing 62. In oneimplementation, the accumulator valve 20 is designed to prevent thevalve housing 62 from exceeding a temperature of 1,000° F.

In one implementation, the area 66 where the distal end of the main bodyinsulation 64 and the main body 48 meet is tapered to reduce the stressproduced when the main body insulation 64 expands due to the heat.Another insulating component or insulating washer 68 is provided justslightly distal of the area 66 to reduce the heat transfer and seal theinterface between the main body 48 and the valve housing 62 at thislocation.

It should be appreciated that the main body 48, valve housing 62, mainbody insulation 64, and insulating component 68 can be made of anysuitable materials. In some implementations, the main body 48 is made ofthe same ceramic matrix composite material as the poppet guide 52 andvalve shaft 54. The valve housing 62 can be made of a light, durablemetal such as titanium.

The insulation 64 can be any suitable material that significantlyinhibits heat transfer from the main body 48 to the valve housing 62. Inone implementation, the insulation 64 is ethylene propylene dienemonomer (M-class) rubber (EPDM). The insulating component 68 can also bemade of any suitable material that significantly inhibits heat transferfrom the main body 48 to the valve housing 62. In one implementation,the insulating component 68 can be made of silica-phenolic material.

As already mentioned, in some implementations, the main body 48, thepoppet guide 52, and the valve shaft 54 are made of a ceramic matrixcomposite material. Any suitable ceramic matrix composite materials canbe used. In one implementation, the main body 48, the poppet guide 52,and the valve shaft 54 are made of carbon zirconium oxide carbide(C—ZrOC) and/or carbon silicon carbide (C—SiC).

Ceramic matrix composites are inherently porous. Those components thatare under pressure, such as the main body 48, may leak hot gas throughthe ceramic matrix composite. In some implementations, the ceramicmatrix composites can be coated with a seal coating 72 (FIG. 15). Forexample, the main body 48 can be coated on the inside and outsidesurface with a seal coating 72. Any suitable material can be used forthe seal coating. In one implementation, the seal coating is a thincoating of silicon carbide (SiC).

Ceramic matrix composite materials are excellent structural insulators.They exhibit structural strength over extreme temperatures while alsoproviding great insulator properties. They also dimensionally stableover a wide temperature range. The ceramic matrix materials with thebest properties for use in the accumulator valve 20 are C—ZrOC andC—SiC.

Ceramic matrix composites are a subgroup of composite materials as wellas a subgroup of technical ceramics. They are made of ceramic fibersembedded in a ceramic matrix, thus forming a ceramic fiber reinforcedceramic material. The matrix and fibers can consist of any ceramicmaterial, whereby carbon and carbon fibers can also be considered aceramic material. In general, the names of ceramic matrix compositesinclude a combination of the type of fiber/type of matrix. For example,C—C stands for carbon fiber reinforced carbon (carbon/carbon), or C—SiCfor carbon fiber reinforced silicon carbide.

Ceramic matrix composites are typically manufactured using the followingthree step process. The first step is to lay-up and fixate the fibersshaped as the desired component. The second step is to infiltrate thefibers with the matrix material. The third step is machining thecomponent and, if required, further treatments like coating orimpregnation of the intrinsic porosity.

The first and the last step are almost the same for all ceramic matrixcomposites: In step one, the fibers, often called rovings, are arrangedand fixed using techniques used in fiber reinforced plastic materials,such as lay-up of fabrics, curtain needled, filament winding, braiding,and knotting. The result of this procedure is called fiber preform orsimply preform.

For the second step, five different procedures can be used alone or incombination with each other to fill the ceramic matrix in between thefibers of the preform: (1) deposition out of a gas mixture, (2)pyrolysis of a pre-ceramic polymer, (3) chemical reaction of elements,(4) sintering at a relatively low temperature in the range 1000-1200°C., and/or (5) electrophoretic deposition of a ceramic powder.Procedures one, two and three find applications with non-oxide ceramicmatrix composites, whereas the fourth is used for oxide ceramic matrixcomposites. It should be appreciated that all of these procedures havesub-variations, which differ in technical details.

The third and final step of machining—e.g., grinding, drilling, lappingor milling—is typically done with diamond tools. Ceramic matrixcomposites can also be processed with a water jet, laser, or ultrasonicmachining.

In some implementations, the main body 48, the poppet guide 52, and thevalve shaft 54 are made using a braided preform. The braided preformprovides greater strength per mass versus other preforms such as curtainneedled preforms. For example, the wall thickness of the main body 48can be reduced by half or more while still maintaining the same pressurerating when a braided preform is used versus a curtain needled preform.

The braided structure provides greater strength because the fibers canbe oriented in the desired manner with minimal cutting. In contrast, thefibers in a curtain needled preform are cut in a Cartesian orientationto fabricate a circular component. Cutting the fibers in this mannerreduces the strength and pressure rating of the resulting ceramic matrixcomposite. In some implementations, the main body 48, the poppet guide52, and the valve shaft 54 can be made of C—ZrOC or C—SiC ceramic matrixcomposites manufactured using a braided preform.

Referring back to FIG. 9, the accumulator valve 20 can be coupled to theaccumulator 16 in such a manner that part of the accumulator valve 20extends into the accumulator 16. This configuration is advantageousbecause it reduces the overall weight and profile of the attitudecontrol system 12.

In some implementations, the main body 48 extends into the accumulator16. When the accumulator 16 is recharged, the main body 48 ispressurized with hot gas. In this state, the main body 48 functions as apressure vessel. When the accumulator 16 is full and the vent valve 22is opened, the pressure inside the main body 48 drops to ambient. Inthis state, the portion of the main body 48 that extends into theaccumulator 16 is under hoop compression by the pressurized gas in theaccumulator 16.

In some implementations, the valve shaft 54 can be held in place at thesecond end 60 of the accumulator valve 20 by a first spacer 76, a secondspacer 78, and a nut 80. The second spacer 78 is coupled to the mainbody 48 using radial pins 82. The nut 80 can be a castle nut thatengages threads on the outside of the second spacer 78. As the nut 80 istightened, it bears down on the valve housing 62 and pulls the secondspacer 78 and main body 48 towards the second end 60 of the accumulatorvalve 20 thereby compressing the insulating component 68.

It should be appreciated that the spacers 76, 78 can be made of anysuitable material. In one implementation, the spacers 76, 78 are made ofan insulating material that inhibits heat transfer to the actuator 38.For example, the spacers 76, 78 can be made of a silica phenolicmaterial and/or a carbo phenolic material.

The accumulator valve 20 includes a throat 84 and a throat retainer 86.The poppet 50 contacts the throat 84 to close the accumulator valve 20.The throat 84 is coupled to the main body 48 at the first end 58 of theaccumulator valve 20. The main body 48 includes a narrow section in thisarea. The throat 84 and throat retainer 86 are positioned on oppositesides of the narrow section of the main body 48 with the throat 84 onthe exterior side and the throat retainer 86 on the interior side. Thethroat retainer 86 is coupled to the throat 84 so that the narrowsection of the main body 48 is sandwiched in between.

It should be appreciated that the throat 84 and the throat retainer 86can be coupled together in any suitable manner. In one implementation,the throat 84 and the throat retainer 86 are coupled together usingthreads. The threads can be oriented in such a way that when the throat84 and the throat retainer 86 are heated, the threads tighten and form aseal that prevents gas from escaping between the throat 84 and main body48.

It should be appreciated that the throat 84 and the throat retainer 86can be made of any suitable materials. In one implementation, the throat84 and the throat retainer 86 can be made of a material that is capableof withstanding high operating temperatures and high velocity gas flows.For example, the throat 84 and the throat retainer 86 can be made ofrhenium molybdenum and/or molybdenum.

Referring to FIG. 15, the interface between the throat 84 and the mainbody 48 is shown. This is one of the areas that can potentially leak ifthese surfaces do not form an adequate seal. One of the difficultieswith this interface is that the throat 84 typically has a much highermodulus than the main body 48, which means the surface of the main body48 will conform to the surface of the throat 84. In one implementation,a slight radius of curvature is provided on the backside of the throat84 to form a corresponding curve on the main body 48. This configurationeffectively seals the interface between these two components.

In some implementations, the accumulator valve 20 satisfies one or moreof the specifications set forth below in Table 6. This design of theaccumulator valve 20 may be especially suitable for use with a low levelattitude control system.

TABLE 6 Accumulator Valve Specifications Parameter Value Contractionratio Min. 3:1 relative to propellant grain Natural throat area Scaledto >1.1x operational throat Permissible leak rate TBD Response time ≥2in/s to 90% full stroke Max. total stroke ≤0.300 in Duty cycle ≥28close/open/close cycles; random operation over 300 s

Vent Valve

Referring to FIG. 16, one implementation of the vent valve 22 is shown.The vent valve 22 includes many of the same components as theaccumulator valve 20. For example, the vent valve 22 includes a poppet88, valve shaft 90, and throat 92. The vent valve 22 moves between anopen position where the poppet 88 is spaced apart from throat 92 and aclosed position where the poppet 88 is in contact with the throat 92. Inone implementation, the actuator 40 moves the valve shaft 90 lengthwiseto move the poppet 88 between the open and closed position.

It should be appreciated that the components in the vent valve 22 can bemade of any suitable material including those already mentioned above inconnection with the accumulator valve 20. For example, the poppet 88 andthe throat 92 can be made of rhenium molybdenum and the valve shaft 90can be made of Inconel 718 or a ceramic matrix composite.

The vent valve 22 can be an off-the-shelf component that is used as isor adapted for use with the attitude control system 12, or it can be acustom designed component. In some implementations, the vent valve 22satisfies one or more of the specifications set forth below in Table 7.This design of the vent valve 22 may be especially suitable for use witha low level attitude control system.

TABLE 7 Vent Valve Specifications Parameter Value Contraction ratio Min.3:1 relative to propellant grain Natural throat area Scaled to ≥1.1xoperational throat Permissible leak rate TBD Response time ≥2 in/s to90% full stroke Max. total stroke ≤0.300 in L* (at max free volume)≥200:1 Pdot rate (at max free ≤10,000 psia/s volume) Duty cycle ≥28close/open/close cycles; random operation over 300 s

Divert Valve

Referring to FIG. 17, one implementation of the divert valve 26 isshown. The divert valve 26 includes many of the same components as theaccumulator valve 20. For example, the divert valve 26 includes a pintle94, pintle guide 96, and throat 98. The divert valve 26 moves between anopen position where the pintle 94 is spaced apart from throat 98 and aclosed position where the pintle 94 is in contact with the throat 98. Inone implementation, the actuator 42 moves the pintle 94 lengthwise intoand out of contact with the throat 98 to close and open the divert valve26.

It should be appreciated that the components in the divert valve 26 canbe made of any suitable material including those already mentioned abovein connection with the accumulator valve 20. For example, the pintle 94and the throat 98 can be made of rhenium molybdenum and the pintle guide96 can be made of Inconel 718 or a ceramic matrix composite.

The divert valve 26 can be an off-the-shelf component that is used as isor adapted for use with the attitude control system 12, or it can be acustom designed component. In some implementations, the divert valve 26satisfies one or more of the specifications set forth below in Table 8and/or Table 9. This design of the divert valve 26 may be especiallysuitable for use with a low level attitude control system.

TABLE 8 Divert Valve Specifications 1 Parameter Value Contraction ratioMin. 3:1 relative to propellant grain Operating throat area 0.00399 in²Natural throat area 0.00439 in² (ø0.075 in) Pintle slope 0.05 in²/inExpansion ratio Max 2:1 relative to operating throat area Permissibleleak rate TBD Response time ≥2 in/s to 90% full stroke Max. total stroke≤0.300 in

TABLE 9 Divert Valve Specifications 2 Parameter Value Required pressurein DACS chamber 500 psia Fill time for ignition ≤2.0 s DACS chambervolume 10 in³ at start 116 in³ at end Divert system ignition valve(DSIV) ballistic requirements Nominal DSIV control pressure 2000 psiaMaximum expected operating pressure 3300 psia Leak rate through DSIVvalve TBD DSIV Sizing requirements Operating throat area (Aero.)0.008659 in² Throat area margin 1.20 Throat area (+margin) 0.010391 in²Natural throat diameter 0.115 in Valve slope 0.06 in²/in Throat geometryTBD Nozzle/exit Maximum allowable divergent half-angle ≤10° (at exit)Maximum exit diameter ≤0.149 in DSIV inlet Inlet tube/barrel flow area≥0.025977 in²

Attitude Thrusters

It should be appreciated that any suitable thrusters 24 can be used withthe attitude control system 12. In general, it is desirable to useattitude thrusters 24 that seal tightly when closed and offerproportional control (versus on/off control). The attitude thrusters 24can provide accurate thruster delivery and minimum impulse bit (MIB)throughout de-pressurization of the accumulator 16.

Operation of the attitude thrusters 24 when the accumulator 16 is notbeing recharged provides the flight vehicle with inherent quiescentthruster delivery that enhances target acquisition capability for flightvehicles such as the KV. The attitude thrusters 24 are preferablylightweight and low cost due to maintaining the gas temperature in theaccumulator <2000° F. enabling uninsulated metallic manifolds andthruster designs. The attitude thrusters 24 can be placed as far aft aspractical to increase pitch/yaw moment capability, which minimizes theattitude control system impulse and thruster levels.

In some implementations, the attitude thrusters 24 satisfy one or moreof the specifications set forth below in Table 10. This design of theattitude thrusters 24 may be especially suitable for use with a lowlevel attitude control system.

TABLE 10 Attitude Thruster Specifications Parameter Value Peak thrust2.5 lbf Thrust rate 125 lbf/s Frequency response 25 Hz operation at ± 1%amplitude and 90° phase Thrust resolution 0.3 lbf Max. impulse (perthruster) 50 lbf · s

Actuators

The actuators 38, 40, 42 can be any suitable actuators. In oneimplementation, one or more of the actuators 38, 40, 42 areoff-the-shelf actuators that are used as it or adapted for use with thevalves 20, 22, 26. In another implementation, the actuators 38, 40, 42are custom designed.

In some implementations, the actuators 38, 40, 42 and/or any otheractuator described in this document satisfy one or more of thespecifications set forth below in Table 11. Actuators satisfying theserequirements may be especially suitable for use with a low levelattitude control system.

TABLE 11 Actuator Specifications Parameter Value Operation typeProportionally commanded Stroke length 0.350 inches (+0.025/−0.000)Operating load 300 lbf, tension and compression (t&c) Min. load vs.position 300 lbf over entire stroke (t&c) profile Inertial load 0.05 lbmMin. slew rate ≥4 in/s over entire stroke and at 300 lbf (t&c) loadingMin. frequency response 25 Hz at ±1% amplitude at −3 dB or 90° phase lagat 300 lbf loading Position accuracy ≤0.002 in over entire stroke and at300 lbf (t&c) loading Position command threshold ≤0.002 in over entirestroke and at 0 and 300 lbf (t&c) loading Duty cycle Continuousoperation for 300+ s at 1 Hz cycling, 100% amplitude, and 100 lbfloading Ambient altitude/pressure Sea-level to high altitude Ambientoperation temp 40° F. to 120° F. Temperature at interface Lineartemperature increase from 75° F. to 300° F. over 300 s

The attitude control system 12 shown in FIG. 18 uses similar componentsand operates similarly to that shown in FIG. 3. The description of thevarious components above applies equally to the same or similarcomponents shown in FIG. 18 unless noted otherwise. Also, thedescription of the operation of the attitude control system 12 appliesequally to that shown in FIG. 18 unless noted otherwise.

One notable aspect of the attitude control system in FIG. 18 is that thevent valves 22 and the divert ignition valves 26 are coupled to theforward end of the gas generators 18. The vent valves 22 and divertignition valves 26 operate in the same or similar manner as describedabove.

One thing to note is the divert ignition valves 26 are coupled to thegas generators 18. This means that hot gas is only available whenpropellant in the gas generators 18 is burning. It should be noted,however, that the divert system 14 can include separate igniters toignite the propellant in the gas generators 28 when hot gas from theattitude control system 12 is unavailable.

In other implementations, the divert ignition valves 26 can be coupledto the accumulator 16 instead of or in addition to being coupled to thegas generators 18. In these configurations, the propellant in the gasgenerators 18 does not need to be burning to provide hot gas to thedivert system 14. However, the accumulator must contain hot gas, whichmeans it must have been initially ignited and/or the gas generators 18initially ignited in order to fill the accumulator 16 with hot gas.

Divert System

Referring to FIG. 18, the divert system 14 includes the divert ignitionvalves 26, the gas generators 28, and the divert thrusters 30. Thedivert system 14 can also include one or more igniters, which can beused to ignite the propellant in the gas generators 28. Any suitableigniters can be used. In some implementations the igniters are barberedpulse igniters.

The gas generators 28 are spaced apart and pneumatically linked by atube 27 (FIGS. 34-36) extending lengthwise between them. The gasgenerators 28 operate together to generate hot gas for the divertthrusters 30. The divert system 14 shown in FIG. 18 includes four divertthrusters 30. Each gas generator 28 is coupled to two of the divertthrusters through openings 350 (FIG. 34) in the side walls.

The nozzles of the divert thrusters 30 can be positioned to exhaustcombustion gases to produce thrust along radial lines separated by90-degree intervals. In some implementations, the nozzles are positionedso that a plane passing through the center of each nozzle also passesthrough, or proximate to, a center of mass of the DACS 10 or the flightvehicle 11.

The divert system 14 can include any suitable number of divert thrusters30 arranged in any suitable configuration. For example, the divertsystem 14 conclude at least 3 divert thrusters, at least 4 divertthrusters, at least 5 divert thrusters, at least 6 divert thrusters, atleast 7 divert thrusters, at least 8 divert thrusters, at least 9 divertthrusters, or at least 10 divert thrusters. The divert thrusters 30 canalso be coupled to the gas generators 28 in any suitable manner. It isgenerally preferable to couple roughly the same number of divertthrusters 30 to each gas generator 28.

The gas generators 28 are positioned opposite each other along thelengthwise axis 36 of the DACS 10 and/or the flight vehicle 11. The gasgenerators 28 are preferably positioned symmetrically on the lengthwiseaxis 36 so that as the propellant burns, it has no effect or only aminimal effect on the overall center of gravity of the DACS 10 and/orthe flight vehicle 11.

The divert ignition valves 26 can have any suitable configuration and bemade of any suitable materials. In some implementations, the divertignition valves 26 can be configured the same or similarly to theaccumulator valves 20 including being made of the same or similarmaterials as the accumulator valves 20. In other implementations, thedivert ignition valves 26 can be configured the same or similarly to anyother valve described in this document or the incorporated documents.FIG. 19 shows a cross sectional view of one implementation of the divertignition valves 26 coupled to the gas generator 18. The vent valve 22 isomitted from FIG. 19.

In some implementations, the divert system 14 can withstand the samepressures and operate for the same amount of time as the DACS 10. Ingeneral, it should be appreciated that any individual parameterdisclosed in connection with the DACS 10 also applies to the divertsystem 14. For example, if the DACS 10 can withstand a given pressure ortemperature, then the divert system 14 can withstand the same pressureor temperature. Also, the operational times of the DACS 10 apply equallyto the divert system 14.

In some implementations, the components in the divert system 14 can bethe same or similar to the components described in connection withsimilar or analogous components in the attitude control system 12. Forexample, the various components in the divert system 14 can be the sameor similar structurally and be made of the same or similar materials asthose in the attitude control system 12.

The divert system 14 can operate in a variety of different ways. In someimplementations, the divert system 14 operates as follows. Thepropellant in the gas generators 28 is ignited by hot gas from theattitude control system 12 or by igniters included as part of the divertsystem 14. The divert thrusters 30 are closed as pressure builds in thegas generators 28.

The pressure in the gas generators 28 builds until it reaches athreshold such as 500 psia. At this point, the divert thrusters 30 areopened to maintain the pressure at about 500 psia. When used to maintainthe pressure, the divert thrusters 30 are operated in a way that doesnot impart a net thrust on the flight vehicle 11. The divert thrusters30 can also be operated to provide a net thrust on the flight vehicle 11in any direction.

Upon completion of the divert maneuvers, the propellant in the gasgenerators 28 is extinguished by simultaneously opening all or enough ofthe divert thrusters 30 thereby causing rapid depressurization andextinguishment of the propellant grain. The divert system 14 enters ahold mode until another divert operation is needed, at which time theprocess is repeated.

The divert system 14 can have numerous configurations. For example, insome implementations, the divert system 14 can be rapidly depressurizedusing vent valves coupled to the gas generators 28.

In general, the divert system 14 provides the lateral motion of theflight vehicle 11, and the attitude control system 12 provides theangular control to stabilize pointing and control divert direction. Thedesign of the attitude control system 12 can be simplified if the centerof gravity of the flight vehicle 11 (e.g., kill vehicle) is aligned withthe divert thrusters 30 and remains aligned throughout operation. Thiscan generally be achieved by positioning some of the avionics componentsaft of the DACS 10. This is called a split configuration as opposed to aunitary layout. The split design may be attractive because the DACS 10is often the highest-risk assembly on the flight vehicle 11.

The divert system 14 can be used to perform a variety of divertmaneuvers, especially in the context of a KV. These include a divertmaneuver using remote sensor data before the operation of the seeker, adivert maneuver after seeker acquisition, a divert maneuver after seekerdiscrimination, and a divert maneuver just before intercept. Of thesedivert maneuvers, the discrimination divert can be the most demandingand can be reduced by the incorporation of a more capable IR seeker. Oneof the main factors when design a KV is the allocation of impulse to thepossible divert maneuvers.

In one implementation, the divert system satisfies one or more of thespecifications in Table 12 below. A divert system meeting theserequirements may be especially suitable for use with the SM-3's KV.

TABLE 12 Divert System Specifications Parameter Value Operation TimeLaunch to initiation: ≤240 s Upon initiation command: ≥1000 s Divertminimum impulse bit <1.8 lbf. s Total divert total delta-V Divert systemdelta-V shall be ≤1900 ft/s Divert initial loaded acceleration Greaterthan or equal to 1.5 G's Divert final acceleration Greater than or equalto 2.0 G's Divert duty cycle ≤4 pulses over 1000 s period Divert pulseoperation time Each pulse shall last ≤4 s Response time (from ignitionResponse time from 0 to 90% thrust command) command from an ignitioncommand shall be ≤2.0 s Response time (when already Response time from 0to 90% thrust ignited) command from a thrust command shall be ≤10 msSatellite Positioning System

FIGS. 48 and 49 show perspective views of one implementation of areaction control system (RCS) 500. The RCS 500 is a spacecraft systemthat uses thrusters to provide attitude control and sometimestranslational movement. The RCS 500 is especially suitable for use withsmall spacecraft such as those used to launch satellites and the like.

The RCS 500 is capable of providing small amounts of thrust in anydesired direction or combination of directions. The RCS 500 is alsocapable of providing torque to allow control of rotation (roll, pitch,and yaw).

The RCS 500 can be used for any of the following: attitude controlduring re-entry, station keeping in orbit, close maneuvering duringdocking procedures, control of orientation, or “pointing the nose” ofthe craft, a backup means of deorbiting, and/or ullage motors to primethe fuel system for a main engine burn.

In some implementations, the RCS 500 is configured to be used withlaunch vehicles such as single-stage sub-orbital launch vehicles andmulti-stage orbital launch vehicles. The launch vehicles can be thosethat only use solid propellant making them much easier to prepare andlaunch. Examples of such launch vehicles include the SPACELOFT andSPYDER systems available from UP Aerospace.

The RCS 500 includes an outer casing or housing 502 enclosing a thrustsystem 504 and an avionics system 506. The housing 502 can have anysuitable configuration that allows the RCS 500 to be coupled to therelevant launch vehicle. The housing 502 shown in FIG. 48 is configuredto be compatible with the payload transport system PTS10 enveloped usedby UP Aerospace. FIG. 50 shows a cross sectional view of the RCS 500coupled to the aft of the nose cone 508 of a launch vehicle. The housing502 can also include various access panels to provide access to thecomponents inside.

The thrust system 504 includes a gas generator 518 pneumatically linkedto thrusters 510. The thrust system 504 also includes igniters 512. Thegas generator 518 includes solid propellant 514 that can be repeatedlyignited and extinguished to provide multiple thrust pulses. The igniters512 are used to ignite the solid propellant 514. The solid propellant514 can be ignited as many times as there are igniters 512.

It should be appreciated that the thrust system 504 can include anysuitable number of igniters 512. In some implementations, the thrustsystem includes at least two igniters 512, at least three igniters 512,or at least four igniters 512.

The solid propellant 514 is extinguished by rapidly depressurizing thegas generator 518. This can be accomplished in a number of ways. In someimplementations, the gas generator 518 is rapidly depressurized by fullyopening all of the thrusters 510. In other implementations, this isaccomplished by opening any suitable number of thrusters 510 anysuitable amount that is sufficient to cause rapid depressurization ofthe gas generator 518.

When the propellant is burning, the thrusters 510 can be used to: (a)maintain the pressure in the gas generator 518 within a range such asthose given in connection with the DACS 10 and/or (b) impart thrust onthe launch vehicle to change its trajectory and/or attitude. Preferably,the propellant is only ignited long enough to operate the thrusters toproduce the desired trajectory/attitude adjustments. This reduces theamount of hot gas that is vented to maintain the pressure with thedesired range in the gas generator 518.

In some implementations, the components in the RCS 500 can be the sameor similar to the components described in connection with similar oranalogous components in the DACS 10 (including the attitude controlsystem 12 and/or the divert system 14). For example, the variouscomponents in the divert system 14 can be the same or similarstructurally and be made of the same or similar materials as those inthe attitude control system 12.

This means the gas generator 518 can be the same or similar to any ofthe gas generators 18, 28, 418. The thrusters 510 can be the same orsimilar to any of the thrusters 24, 30, 430. These components can alsobe the same or similar to any similar component that is described in theincorporated documents. For example, the thrusters 510 can be the sameor similar to those described in the incorporated document titled “HotGas Thruster.” Nearest other examples are also possible.

The avionics system 506 includes those components used to control andoperate the RCS 500. In some implementations, this includes actuationand actuation control electronics, control processor, data acquisitionI/O, power regulation, power supply (battery), signal conditioning,embedded software, and cable and harnessing.

In some implementations, the igniter 512 can be configured in the mannershown in FIGS. 54-56. The igniter 512 includes a housing 519, apropellant grain 520, a barrier support disk 522, a barrier 524, athroat 526. Insulated copper wire 528 extends through the housing 519and is connected to exposed nichrome wire 532. The nichrome wire 532 ispositioned adjacent to the propellant grain 520. When a voltage isapplied to the wire 528, it heats the nichrome wire 532, which ignitesthe propellant grain 520. The hot gas produced by the burning propellantbursts the disk and enters the gas generator 518 igniting the propellant514. The igniter 512 can be sized to light at final free volumes withoutover-pressurizing at initial free volumes.

FIGS. 58 and 59 show perspective views of another implementation of anRCS 600 configured to fit in SpaceLoft's avionics envelope thatcurrently houses its cold gas system. The RCS 600 includes a housing602, the thrust system 504, and an avionics system 606. FIG. 58 showsthe thrust system 504 and the avionics system 606 in the housing 602.FIG. 59 shows the RCS 600 with the housing 602 removed.

FIGS. 60-63 illustrate the operation of the thrust system 504. FIG. 60shows the initiation of the thrust system 504 from an off state. Theigniter 513 is ignited (the other igniters 512 have not been ignited),bursts the barrier 524, and ignites the propellant grain 514 in the gasgenerator 518. The proportional thrusters 510 take over and the thrustsystem 504 transitions into thrust mode. The thrusters 510 are closed inFIG. 60.

FIG. 61 shows the thrust mode of the thrust system 504. The thrusters510 maintain the pressure at a given level such as 500 psia. Thethrusters direct the hot gas to deliver a given amount of thrust (e.g.,9 lbf) in any direction. The nominal thrust duration can be any suitableperiod of time such as 4 seconds. The igniter barriers 524 prevent theigniter propellant grain 520 in the unspent igniters 512 from igniting.The thrust system 504 transitions to extinguishment mode.

FIG. 62 shows the extinguishment mode of the thrust system 504. All ofthe thrusters 510 are commanded fully open. The large change in throatarea and rapid depressurization rate (the rate can be the same as thatdisclosed for the DACS 10) extinguishes the propellant grain 514. Thethrust system 504 transitions into hold mode.

FIG. 63 shows the hold mode of the thrust system 504. During hold times,the thrusters 510 remain open. The RCS 500, 600 enters low power, idlemode. Health and functional checks may be performed in hold mode. Thethrust system 504 transitions into initiation mode upon command from thelaunch vehicle.

In some implementations, the various RCS 500, 600 satisfy one or more ofthe specifications set forth below in Table 13. This design of the RCS500, 600 may be especially suitable for use with small spacecraft usedto launch satellites.

TABLE 13 Reaction Control System Specifications Parameter Value Totalweight 18 lbf Packaged impulse 230 lbf · s Total burn time 18 s Pulses4x Deliverable system thrust 13 lbf · s Avionics operation time Durationof mission Delivered thrust pulse time ≤8 seconds per pulse ≤4 discretepulses Initiation delay <1.5 sec to 90% design thrust Initiator charge,final 450 ± 50 psia delivered Initiator barriers Burst at 600 ± 50 psia.Seal from manifold side up to MEOP Generated RCS thrust 12 lbf DeliveredRCS thrust <2% magnitude error Delivered RCS thrust <2° misalignment RCSMEOP 550 psia GG dimensions OD: 4.49 in L: 1.73 in Mechanical interface21804-00180-01 Electrical interface 21804-00184-01 Software interface21804-00188-01

EXAMPLES

The following examples are provided to further illustrate the disclosedsubject matter. They should not be used to constrict or limit the scopeof the claims in any way.

Example 1

A hot fire test of a hot gas attitude control system 150 was performedusing the prototype system shown in FIGS. 20-21. The prototype system150 was used to demonstrate the feasibility of such a system when usedas part of a solid propellant DACS (SDACS) for a guided missile. The hotgas attitude control system would provide hot gas to: (1) the thrustersthat control the attitude of the guided missile and (2) the propellantin the divert system to ignite one or more times as part of a divertoperation.

It should be noted that the prototype system 150 is not identical to asystem that would be used on a guided missile. However, the components,internal materials, ballistic configuration and envelope of theprototype system 150 are representative of a flight design. Thus, theprototype hardware and associated hot fire test results can be used toassess the feasibility of flight ready low level attitude control systemdesign such as the one shown above.

The prototype system 150 included an accumulator 152, a gas generator(GG) 154, an accumulator valve 156, a vent valve or extinguishment valve158, an expansion port 160, and an accumulator valve housing assembly162. The prototype system 150 also included an accumulator valveactuator (not shown) and a vent valve actuator 166. The actuators areconventional actuators used in these types of applications. Theprototype system 150 included various sensors (not shown) to collectimportant operational characteristics such as pressure and temperature.

As shown in FIGS. 20-21, the gas generator 154, the accumulator valve156, the vent valve 158, and the expansion port 160 were all operativelycoupled to the accumulator valve housing assembly 162. The accumulatorvalve housing assembly 162 included a central passage 164 through whichhot gas can flow between each of the attached components. Theaccumulator valve 156 was positioned between the accumulator 152 and thepassage 164 to control the flow of hot gas to and/or from theaccumulator 152.

The prototype system 150 was set up as follows. A start propellant grainwas positioned in the accumulator 152 with the rest of the propellantbeing in the gas generator 154. The expansion port 160 was capped with aburst disk. The expansion port 160 was included so that a divert systemcan be coupled to the system 150 in future tests. In such aconfiguration, a divert system ignition valve would be coupled to theexpansion port 160 to selectively and repeatedly allow hot gas into thedivert system to ignite the propellant for divert operations.

FIGS. 22-27 and Table 14 to Table 19 show the structure and materialsfor the accumulator 152 (FIG. 26; Table 18), gas generator 154 (FIG. 27;Table 19), accumulator valve 156 (FIGS. 22-23; Table 15), vent valve(FIG. 24; Table 16), and the accumulator valve housing assembly 162(FIG. 25; Table 17).

TABLE 14 Description of Materials Material Description Moly MolybdenumReMo Rhenium molybdenum 17-4 H1150 17-4 H1150 stainless steel alloyC-ZrOC Carbon zirconium oxide carbide ceramic matric compositeS-phenolic Silica phenolic C-phenolic Carbon phenolic EPDM Ethylenepropylene diene monomer (M-class) rubber Inconel 718 Nickel chromiumalloy 300 Series 300 series austenitic stainless steel GaroliteReinforced phenolic material Garolite CE Medium weave cotton clothphenolic

TABLE 15 Accumulator Valve Materials (FIGS. 22-23) Ref. No. NameMaterial 200 Poppet guide Moly 202 Poppet ReMo 204 Housing 17-4 H1150206 Valve body C-ZrOC 208 Conic seal S-phenolic 210 Valve body insulatorEPDM 212 Throat retainer Moly 214 Throat ReMo 216 Shaft shield Moly 218Standoff insulator C-ZrOC 220 Accumulator shaft C-ZrOC 222 Actuatorclosure 17-4 H1150 224 Actuator adapter Inconel 718 226 Retaining pinTungsten 228 Retainer nut Inconel 718 230 Retainer insulator S-phenolic232 Collar insulator S-phenolic 234 Collar retainer Inconel 718 236Wavespring

TABLE 16 Vent Valve Materials (FIG. 24) Ref. No. Name Material 238 Ventvalve body C-ZrOC 240 Vent plenum insulator S-phenolic 242 Vent poppetReMo 244 Vent shaft C-ZrOC 246 Vent actuator adapter Inconel 718 248Vent throat ReMo 250 Vent seal closure 17-4 H1150 252 Vent valve bodyEPDM insulator 254 Vacuum tube S-phenolic insulator 256 Wavespring

TABLE 17 Accumulator Valve Housing Assembly Materials (FIG. 25) Ref. No.Name Material 258 GG inlet insulator S-phenolic 260 Gas tube insulatorS-phenolic 262 GG castle nut 300 series 264 Burst disk insulatorS-phenolic 266 Burst disk closure 17-4 H1150 268 Centering housing 300series 270 Centering bullet 300 series 272 Centering shaft 300 series274 Centering bracket 300 series 276 Actuator bracket Aluminum 278Actuator base Aluminum 280 Accumulator castle 300 series nut

TABLE 18 Accumulator Materials (FIG. 26) Ref. No. Name Material 282Accumulator closure 17-4 H1150 284 End cap 17-4 H1150 286 Bleed orificeC-phenolic insulator 288 Orifice entrance C-phenolic insulator 290 Bleedorifice Moly 292 Accumulator 17-4 H1150 chamber 294 Case sleeveinsulator EPDM assembly 296 Case sleeve Garolite 298 Front plateinsulator EPDM assembly 300 Front plate Garolite CE 302 Rear plateinsulator EPDM assembly 304 Rear plate Garolite

TABLE 19 Gas Generator Materials (FIG. 27) Ref. No. Name Material 306 GGchamber 17-4 H1150 308 End cap 17-4 H1150 310 GG closure 17-4 H1150 312GG forward insulator C-phenolic 314 GG rear shim Garolite CE insulator316 GG tuber spacer S-phenolic insulator 318 Propellant cup AAP-3797 320GG propellant Garolite CE base 322 GG propellant Garolite CE sleeve 324Propellant cup EPDM insulator sleeve 326 Propellant cup EPDM insulatorbase

The accumulator valve body 206 had a 0.300 inch wall thickness and wasdesigned to withstand a maximum expected operating pressure of 2,250psia and a maximum operating temperature of 2,000° F. The othercomponents in the prototype system 150 were designed to withstand amaximum expected operating pressure of 3,500 psia. This meant that theC—ZrOC components drove the design of the other structures.

The accumulator valve body 206 and the vent valve body 238 were coatedwith 0.0010±0.0005 inch of silicon carbide (SiC) to prevent hot gas fromflowing through these components. The valve bodies 206, 238 are made ofC—ZrOC, which is inherently porous. Hot gas can leak through these partswhen they are pressurized. The SiC coating helps prevent hot gas fromleaking. Also, the hot fire tests revealed that the particles in the hotgas also help to plug and seal the pores in the C—ZrOC components.

The hot fire test had the following primary objectives: (1) demonstrateoperation of the accumulator valve 156 for 200 s, (2) demonstrateoperation of the accumulator valve poppet 202, control system, and gasflow operations, and (3) demonstrate basic propellant operationsincluding ignition, extinguishment, and reignition. The hot fire testhad the following secondary objectives: (1) demonstrate basicballistics, (2) measure burnback of the propellant, pressure drops, andperformance of the accumulator 152, (3) demonstrate control logic, and(4) demonstrate rack operation, vacuum, and ignition system.

The prototype system 150 was configured to operate in the followingmanner. An initial start propellant grain is ignited in the accumulator152 with the accumulator valve 156 closed. The pressure rises in theaccumulator until it exceeds 1,260 psia. At this point, the controllerinitiates a recharge event by opening the accumulator valve 156 andallowing hot gas to enter the gas generator 154 and ignite thepropellant 318.

The hot gas flows from the gas generator 154 to the accumulator 152until it reaches a pressure of 1900 psia. The gas generator 154 isextinguished by closing the accumulator valve 156 and opening the ventvalve 158. The pressure drops in the accumulator 152 as hot gas exitsthrough the bleed orifice 290. Another recharge event is initiated whenone of the following events occurs: (1) the pressure reaches a minimumlevel in the accumulator 152 or (2) ten seconds have elapsed. Theminimum pressure level in the accumulator 152 was set at 1,000 psia forthe first three recharges and 500 psia thereafter. The hot fire test isconducted under conditions that simulate high altitude >50,000 ft andtemperatures of 40-90° F.

Before the hot fire test, the prototype system 150 was pressure and leaktested using inert gas. The accumulator valve 156 was tested to verifythat it moved accurately and without issues. The other hardware in theprototype system 150 was tested to verify that its performance wasacceptable for the purposes of the test. The propellant 318 was X-rayedto ensure no cracks or voids existed in the grains which could causeunintended consequences during a test. The prototype system 150 wassecured inside a modified magazine.

FIG. 28 shows the test data in its completeness and, for all intent andpurposes, indicates no major anomalies. The initial pressurizationcharge in the accumulator 152 successfully triggered the softwarecontroller and started a series of recharges. The first three rechargesare pressure triggered when the accumulator reaches approximately 1,400psia. The remaining recharges occurred after the ten second timeoutperiod elapsed. In total, twenty one recharges occurred in the specified200 second mission time, and afterward pressure in the accumulator washeld for an addition 300 seconds.

The pressure in the prototype system 150 stayed well below the maximumexpected operating pressure and peaked at 1,941 psia. FIG. 29 shows adetailed record of the initial pressurization and first recharge. Themajor events are denoted by vertical lines A though H and described asfollows.

Event A in FIG. 29 denotes the ignition of the accumulator charge andinitial pressurization of the accumulator. At T-0 seconds, power wasapplied to the nichrome wire to initiate heating of the accumulatorpressurization propellant. It took approximately 3.0 seconds for thewire to reach a critical temperature and ignite the initial propellantcharge. Within 0.25 seconds, the pressure in the accumulator 152 rapidlyincreased thereby indicating that the propellant charge was fullyignited.

Event B occurred at T+3.40 seconds. At this point, the pressure in theaccumulator 152 exceeded the 1,260 psia threshold and activated the testcontroller. The test was now running in closed-loop operation.Simultaneously, the controller initiated a recharge event and commandedthe accumulator valve 156 to open at the specified 0.5 in/s slew rate.

Event C occurred at T+3.64 seconds. At this point, the accumulator valve156 reached a critical position, approximately 0.055 inches, and hot gasbackflowed from the accumulator 152 to the gas generator 154. Theaccumulator shaft 220 deflected a small amount due to the increasedpressure load. Within 0.060 seconds, the pressure in the accumulator 152and the gas generator 154 equalized and the burning propellant 318 beganto increase the pressure in the accumulator 152.

Event D occurred at T+4.04 seconds. At this point, the pressure in theaccumulator 152 reaches the 1,900 psia trigger. The controller commandedthe vent valve 158 to begin opening. By T+4.24 the pressure in the gasgenerator 154 dropped back to ambient and the pressure in theaccumulator 152 sealed the poppet 202 closed. For the next severalseconds the pressure in the accumulator 152 was steadily exhaustedthrough the bleed orifice 290.

Event E occurred at T+5.84 seconds. At this point, the pressure in theaccumulator 152 reached 1,400 psia and the vent valve 158 started toclose to initiate a recharge event. It should be noted that the 1,400psia limit was intentionally set higher than the 1,000 psia desiredrecharge pressure so that the vent valve 158 was closed for a shortamount of time to determine if the propellant 318 was smoldering. Thepressure in the gas generator 154 between event E and F remained steadyat approximately 0 psia, meaning the grain was fully extinguished.

Event F occurred at T+6.08 seconds. At this point, a 0.25 second timeoutoccurs and the accumulator valve 156 was forced to start opening eventhough the pressure in the accumulator 152 is well above 1,000 psia at1,380 psia. This was partially due to an inaccurate bleed-down rate—thepressure was expected to have dropped significantly more due to heattransfer to the walls and mass loss through the bleed orifice 290. Audiorecording obtained as part of the test data revealed a periodic“whistling” from the bleed orifice 290 that fluctuated in intensity andindicated a partial clog. This partially explained why the pressure didnot drop as fast as predicted.

Event G occurred at T+6.32 seconds. At this point, the accumulatorpoppet 202 opened to the critical position and allowed hot gas from theaccumulator 152 to backflow into the gas generator 154 to initiate arecharge.

Event H occurred at T+6.88 seconds. At this point, the accumulatorreached the 1,900 psia trigger and the process of extinguishing the gasgenerator 154 began. From here, the general pattern repeated itselfsuccessively. It should be noted that clogging of the bleed orifice 290became more evident as the test continued. FIG. 29 shows that thepressure in the accumulator 152 at recharge slowly increased fromapproximately 1,000 psia up to 1,300 psia by the end of the test. Theaudio recording also confirmed that the “whistling” from the bleedorifice 290 was not as audible.

The hot fire test completely fulfilled all of the primary and secondarytest objectives. The performance of the actuator for the accumulatorvalve 156 was in line with expectations and the control algorithm keptthe pressure in the accumulator 152 below the maximum expected operatingpressure. The clog in the bleed orifice 290 caused ten second timeoutsand recharges for the majority of the test. Because of this, thepressure in the accumulator 152 never dropped below the 500 psiathreshold.

Example 2

A hot fire test of the hot gas attitude control system 150 was performedusing the prototype system shown in FIGS. 20-21 with a modified dutycycle. The goal of this test was to extend the duty cycle to 300+seconds by increasing the time between recharges. The prototype system150 was largely the same as in Example 1 except that some of the sensorsand instrumentation were upgraded. Prior to running the test, thehardware was tested using the same procedures described above in Example1.

FIG. 30 shows the results of hot fire test. This test did not meet itsprimary objective of demonstrating multiple recharges in a 300 secondduty cycle. As shown in FIG. 30, the initial propellant chargepressurized the accumulator 152 to approximately 1,350 psia andactivated the controller at approximately T+2 seconds. The pressure inthe accumulator 152 dropped to 1,125 psia at approximately T+3 secondsand initiated a recharge sequence (i.e., accumulator valve 156 wasopened) that ignited the propellant 318 in the gas generator 154. Thepropellant 318 in the gas generator burned until the pressure in theaccumulator 152 reached 1,975 psia at approximately T+4 seconds. Thepressure in the accumulator 152 was allowed to bleed down forapproximately 10 seconds to 1,125 psia when another recharge sequencestarted.

FIG. 30 shows that the pressure in the gas generator 154 rapidly reachedequilibrium with the pressure in the accumulator 152 but there is noindication that the propellant 318 reignited. The accumulator valve 156remained open and the test continued for approximately 380 secondswithout the propellant 318 reigniting.

After evaluating the thermal test date, the cause of the reignitionfailure is believed to be the ten second dwell time between the lastignition and the subsequent ignition attempt. The prototype system 150has a large thermal mass that absorbed too much of the heat between thefirst ignition event and the subsequent failed reignition attempt. Thetemperature of the hot gas was too low at the time of the failedreignition event to ignite the propellant 318.

Despite the failed reignition, the control logic continued to operatenominally. The controller recognized that it failed to ignite andcontinued to command a recharge until the test was manually stopped.

Example 3

A hot fire test of the hot gas attitude control system 150 was performedusing the prototype system shown in FIGS. 20-21 to correct the problemsidentified in Example 2 and extend the duty cycle to 500+ seconds. Theprototype system 150 was largely the same as in Example 2. Prior torunning the test, the hardware was tested using the same proceduresdescribed above in Example 1.

The duty cycle was modified in the following ways based on the test inExample 2. The pressure level at which the accumulator 152 would triggera recharge was changed from 1,125 psia back to 1,400 psia (what it wasin Example 1). The duty cycle was modified to include a warm-up periodwhere the first three recharges are subject to a 2.5, 3.0, and 3.5second timeout. What this means is that the first recharge would beinitiated after 2.5 seconds, the second after 3.0 seconds, and the thirdafter 3.5 seconds regardless whether the pressure in the accumulator 152had dropped below the low pressure level.

After the warm-up period the duty cycle was set to revert to a tensecond timeout for the first minute of the test. After the first minute,the recharge timeouts were gradually increased from 10 seconds to 25seconds. After 325 seconds, the control logic transitioned into anextended mission mode where the recharge timeout and minimum pressurerecharge trigger were set aggressively to 45 seconds and 750 psia,respectively. The test was set to run indefinitely until it was manuallystopped.

The results of the hot fire test are shown in FIG. 31. The modificationsto the duty cycle successfully extended the operational time to 500+seconds. The first sixty seconds of the duty cycle in this test matchclosely the same data from the test in Example 1. This duty cycle does agood job of thermally conditioning the system 150 as shown by the factthat all subsequent recharges occurred without incident.

At T+325 seconds the test successfully demonstrated the capability ofperforming four recharges 45 seconds apart before running out of thepropellant 318 midway through a recharge at T+500 seconds. When thepropellant rant out, the accumulator valve 156 was retracted and the hotgas inside the accumulator 152 was held for an additional 400 secondsfor a total mission time of 900 seconds. During this time, the pressurein the accumulator 152 gradually decayed at an average rate ofapproximately −3.25 psi/s due to a partial clog in the bleed orifice290. At T+905, the test was stopped and the pressure was vented from thesystem 150. It should be noted that the peak pressure during all therecharges stayed below 1,986 psia, which is only slightly above thetarget pressure of 1,975 psia and well below the 2,500 psia maximumexpected operating pressure.

This test consumed the entire propellant grain in an effort todemonstrate the maximum capability of the system 150. Assuming atargeted flight-I_(sp) of 185 sec, the full 1.1 lbm of propellant 318 isequivalent to 204 lbf·s of impulse through one valve. This is asubstantial improvement over the target amount of only 100 lbf·s ofimpulse over 300 seconds of operation.

The hot fire test satisfied all primary and secondary objectives. Thesystem 150 demonstrated 24 recharges spanning a 325 second time frame byrevising the initial duty cycle to match the previously successful testin Example 1. Afterward, the system 150 used aggressive rechargetimeouts and pressure triggers to demonstrate an additional fourrecharges with 45 second dwell times. The current system 150 andespecially the accumulator valve 156 show that it has a substantialmargin for error. This shows that there is an opportunity tosignificantly reduce the weight of the system 150 and/or implement dutycycles well in excess of 500 seconds and 200 lbf·s of impulse through asingle valve.

Example 4

A prototype divert system 400 is designed and fabricated. Isometricviews from the forward and aft directions of the divert system 400 areshown in FIGS. 32 and 33. The divert system 400 is a closerepresentation of a production divert system such as the one shown inFIG. 18. The divert system 400 is designed to satisfy the requirementsin the following tables.

TABLE 20 Prototype Divert System Specifications Parameter ValuePropellant loading 2.5 in usable web. Outer diameter driven by maxenvelope and linings. Operation time Launch to initiation: ≤1000 s Uponinitiation command: ≥1000 s Divert duty cycle Requirement: 3 pulses over1000 s period (POD at T+0, T+500, T+1000) + 1 backup pulse/igniterDivert pulse 1 s operation time Divert operating 80 lbf nominal pressureDivert system Not specified thrust Maximum The maximum mass, when fullyloaded with mass expendables, shall not exceed 300.0 lbm Maximum 17.26in diameter × 225 in length envelope Thruster TBD misalignmentStructural Maximize as best as possible within reasonable stiffnessjudgement. Normal temp System shall meet performance and reliabilityafter continuous exposure to assembly and check out temperature range of+50° F. to +95° F. during the final assembly and missileencanisterization period of up to 180 days. Air conditioning Systemshall meet performance and reliability malfunction temp after continuousexposure to assembly and check out temperature range of 0° F. to +120°F. due to air conditioning malfunction for a period not to exceed 72hours. There can be up to three air conditioning malfunctions during any180 day period. Storage temp System shall meet performance andreliability after continuous exposure to storage temperature range of−20° F. to +120° F. for periods up to 2 years. The conditions apply withmaximum variation in any 24 hour period of 30° F. Transportation Systemshall meet performance and reliability temp after exposure totransportation temperature range of −20° F. to +130° F. for periods ofup to 5 days. Normal pressure System shall meet performance andreliability after exposure to assembly and checkout atmosphericpressures of 15.4-11.3 psia (sea-level to 7000 ft). PHS&T pressureSystem shall meet performance and reliability after exposure to PHS&Tatmospheric pressures of 15.4- 2.7 psia (sea-level to 40,000 ft).Factors of safety The factors of safety for structures shall be asfollows: All structures Not Specified: Yield: 1.50 Ultimate: 2.00Fittings (Ultimate): 1.75 Castings (Ultimate): 2.00 DACS PressurizedComponents: Proof: 1.10 × MEOP Yield: 1.50 Burst: 2.00 × MEOP DACSPropellant Grain and Bondline: Ultimate: 2.00 DACS Pressurant Tank:Proof: 1.50 × MEOP Burst: 2.00 × MEOP Margin of safety System shall havepositive strength margins of safety based on minimum material thickness,minimum strength, extreme temperatures, and maximum expected combinedloads calculated as follows for yield and ultimate conditions: MS =X/(Y*FS) − 1 > 0.0

TABLE 21 Prototype Attitude Control System Specifications ParameterValue Operation time Launch to initiation: ≤1000 s Upon initiationcommand: ≥1000 s Total ACS impulse Not to exceed 50 lbf s per thrusterACS duty cycle On when divert thrusters are on ACS pulse operation timeContinuous operation while divert thrusters are on

TABLE 22 Prototype Divert Thruster Specifications Parameter ValueNominal peak thrust 48.09 lbf 80.72 lbf overthrust capable at 3000 psiaThrust accuracy Better than or equal to 0.30 lbf (at 1365.21 psia)Nominal operating pressure 1365.21 psia 3000 psia during overthrustMaximum expected operating 3300 psia pressure (MEOP) Throat Operatingthroat area (Aero.) 0.021410 in² Throat area margin 1.20 Throat area(+Margin) 0.025692 in² Natural throat diameter 0.1809 in Valve slope0.11 in²/in Throat geometry 0.47 in²/in Class Circle-in-circle dBarrel1.250 in thetaInlet 45° dThroat 0.388 in rsi 0.125 in thetaExit 15°dExit 1.972 in rTip 1.500 in yTipStart 1.160 in dShaft 0.500 in dSeal0.343in Nozzle Maximum allowable divergent ≤15° half-angle (at exit)Minimum exit diameter ≥0.590 in Thruster inlet Inlet tube/barrel flowarea ≥0.06423 in²

TABLE 23 Prototype Divert Thruster Actuator Specifications ParameterValue Step Step constant resistance force ≥75 lbf 0-90% step timeagainst constant force ≤0.010 s Step size 0.195 in Bandwidth Forcebandwidth at −3dB, −90° ≥75 Hz Bandwidth sinusoidal half amplitude 0.001in input (±in) Constant prevailing torque driven during 75 lbf bandwidthDuty cycle Duty cycle RMS force 75 lbf Duty cycle time 16 s Temp riseduring duty cycle ≤150° F. Threshold High load threshold 0.001 in Loadat which threshold is measured 75 lb Accuracy Accuracy (±in) ≤0.001 inTemperature range for accuracy 40-350° F. Load for accuracy measurement75 lb Miscellaneous Soak temperature Not specified Actuator weight ≤0.75lb Inertial load 0.015 lb Friction load 12.5 lb Minimum total travelReq: 0.245 in (+.025/−.000) Goal: ≥0.345 in No load speed (m/s) Notspecified Loaded speed (in/sec) >20 in/s to meet 0-90% step at 75 HzStall force ≥75 lbf Position sensor DVRT, LVDT, potentiometer, orsimilar Supply voltage [V] 24-33 VDC (regulated) Max steady state powerdraw ≤150 W Position command signal 0-10 VDC, 25 mA max Remoteenable/disable 5 V TTL enable/disable Current draw sense Linearly scaled0-10 VDC current feedback Position Feedback Linearly scaled 0-10 VDCposition feedback on output shaft

TABLE 24 Prototype Attitude Thruster Specifications Parameter ValueNominal thrust 4.08-16.32 lbf (500-2000 psia) Thrust accuracy Betterthan or equal to 0.20 lbf (500-2000 psia) Nominal operating pressure500-2000 psia Maximum expected operating 2200 psia pressure Allowableleak rate through ACS TBD thruster Pneumatic sealing load at 500 psiaTBD, forcing shaft into throat Throat Operating throat area (Aero.)0.004902 in² Throat area margin 1.20 Throat area (+margin) 0.005883 in²Natural throat diameter 0.0865 in Valve slope 0.06 in²/in Throatgeometry 0.47 in²/in Class Circle-in-circle dBarrel 1.250 in thetaInlet450 dThroat 0.388 in rsi 0.125 in thetaExit 15° dExit 1.972 in rTip1.500 in yTipStart 1.160 in dShaft 0.500 in dSeal 0.343in Nozzle Maximumallowable divergent ≤18° half-angle (at exit) Minimum exit diameter≥0.282 in Thruster inlet Inlet tube/barrel flow area ≥0.014706 in²

TABLE 25 Prototype Attitude Thruster Actuator Specifications ParameterValue Step Step constant resistance force ≤10 lbf 0-90% step timeagainst constant ≥0.003 s force Step size 0.082 in Bandwidth Forcebandwidth at −3 dB, −90° ≥226 Hz Bandwidth sinusoidal half amplitude0.001 in input [±in] Constant prevailing torque driven 10 lbf duringbandwidth Duty cycle Duty cycle RMS force 2.5 lbf Duty cycle time 1000 sTemp rise during duty cycle ≤150° F. Threshold High load threshold 0.001in Load at which threshold is measured 10 lb Accuracy Accuracy [±in]≤0.001 in Temperature range for accuracy 40-350° F. Load for accuracymeasurement 10 lb Miscellaneous Soak temperature Not specified ParameterValue Actuator weight ≤0.50 lb Inertial load 0.010 lb Friction load 12.5lb Minimum total travel 0.120 in (+.025/−.000) No load speed (m/s) Notspecified Loaded speed (m/s) >26 in/s to meet 0-90% step at 226 Hz Stallforce ≥25 lbf Position sensor DVRT, LVDT, potentiometer, or similarSupply voltage 24-33 VDC (regulated) Max steady state power draw ≤150 WPosition command signal 0-10VDC, 25 mA max Remote enable/disable 5V TTLenable/disable Current draw sense Linearly scaled 0-10VDC currentfeedback Position feedback Linearly scaled 0-10VDC position feedback onoutput shaft

The divert system 400 includes four divert thrusters 430 coupled tocorresponding electromechanical (EM) actuators 402 positioned betweentwo middle domes or thruster interface domes 404. Four barrier ignitors406 are coupled to the domes 404—two on each dome—to provide four divertignition pulsing events. The ignitors 406 are pneumatically linked tothe gas generators 418 through ports 408. Other ports 410 are used tocouple thermocouple and pressure transducer instruments to the domes404. The divert system 400 includes two extinguishable solid propellantgas generators (GG) 418 that mirror each other and operate (ignite, burnand extinguish) simultaneously.

An insulated hot gas tube 427 is centered on the axis of the gasgenerators 418 and pneumatically links the gas generators 418 together.The tube 427 provides pressure commutation and structure. There is askirt 412 at the outermost end of each gas generator 418. The skirt 412is used to mount hardware to the ground test table, rocket bulkheadinterface and any other accessories that might be required. Within eachgas generator 418 is an extinguishable solid propellant grain 414. Theoverall dimensions of the divert system 400 are about 5 inches indiameter and 15 inches in length.

The components and materials of the gas generators 418 are shown inFIGS. 37-42. The inner dome 404 contains a number of features. Each dome404 has an interface 416 to pneumatically link the gas generators 418 toeach other and balance pressure between each side.

Each gas generator 418 is coupled to two divert thrusters 430 for atotal of four divert thrusters 430 (proportional divert thrusters).Metallic gas generators 418 were used for the prototype divert system400. Other gas generator 418 materials and manufacturing options can beused in production systems such as insulated composites wound andcomposite overwrap of metallic liner.

Two different sizes of ports 408, 410 are used. The smaller ports 410are used for instrumentation and for adapting to the divert ignitionvalve. The larger size ports 408 are used for igniter interfaces. Thelarge threaded interface 416 between the dome and a case or housing 420allows cartridge loading of the propellant grains (production systemswill likely cast propellant directly into cases to minimize overallenvelope and weight). Each gas generator 418 contains half of the totaldivert subsystem propellant.

There is a large fiber-reinforced phenolic spacer 422 at the outer endof each gas generator 418. The spacer 422 is meant to take up the domevolume so that the propellant grain 414 is flat on both ends. A standardelastomeric O-ring bore seal is used to contain pressure inside the gasgenerator 418. The spacer 422 is included in the prototype only for testefficiency in using existing hardware designs. A production system willfill the entire available volume of the gas generator 418 withpropellant 414 to maximize overall propulsion system mass fraction.

FIG. 40 illustrates details of the interface between the dome 404 andthe case 420 of the gas generator 418. One challenge of the design is toaccommodate a full diameter case joint by capturing the dome 404 whileusing the case 420 to limit dome deflection under pressure. The casejoint incorporates buttress threads which prevent the thin walledthreaded ring from unseating due to ejection loads. The O-ring sitswithin the threaded interface and capture link. Extrusion gaps aremaintained as a result of this interface design.

FIG. 41 shows additional details of the pneumatic link interface betweenthe gas generators 418. A spiral retaining ring 424 is loaded with acastle nut 426 and thereby mates the two gas generators 418 to eachother. The linking tube 427 is sealed using an O-ring bore seal 428. Allof the internal surfaces are insulated to limit hardware temperaturesand preserve energy in the gas.

FIGS. 43-45 show additional details of the divert thrusters 430. Thedivert thruster 430 includes a main bracket 432 that turns the actuatoraround 180° to allow more room for the actuator motor. Turning theactuator induced large relative displacements (and therefore thrustinaccuracy) between the actuator and thruster pintle. In order tominimize the relative displacements and deflections, the actuator linkand slide carriage 434 were designed. They add substantial stiffness tothe system. The orientation of the two items also maintainsproportionality between the thruster pintle position and actuatorstroke. The slide and carriage assembly 434 further improve bendingstiffness.

Attitude thrusters 436 are added to the prototype divert system 400 todemonstrate the attitude thruster technology in parallel with the divertprototype 400. The integral ACS manifold is shown in FIG. 46. Thepassageway 438 is machined directly into the outside of the gasgenerator case 420. Two EPDM insulators 440 are assembled in thepassageway 438 and bonded into place. The attitude thruster 436 isassembled into place with a male bore seal with a clocking tab tocontrol orientation. The attitude thruster 436 is held in place with ashoulder retaining clip.

Like the ACS manifold, the igniter 406 is integral with the gasgenerator case 420. The throat, burst disk, igniter propellant, andcushion are assembled into the case 420 in that order. A closure with aPacSci initiator caps off the igniter 406. It is sealed with a male boreseal held in place with a retaining ring. The igniter gas directlyenters the gas generator 418 through the carbon phenolic throat.Subsequent igniters 406 are prevented from premature ignition by theone-way burst disks.

Example 5

The thermal heat transfer associated with a 1000 s propulsion duty cycleis analyzed. The heat transfer drives the overall dimensions and designtechniques of the gas generator chamber and associated components. Theheat transfer coefficients change dramatically depending on the locationwithin the gas generator 418. The thruster inlets see the highest heatload and duration. The main propellant grain is not modeled in thethermal analysis. Instead, the heat is applied to the walls of the gasgenerator 418 in a burn-back duty cycle to simulate the heating into thecase 420 and the insulation as the propellant grain 414 regresses andexposes the gas generator 418 to hot gas.

The location of each convective thermal load has three main event types:ignition, pulse, and off. This analysis does not include any radiationand is adiabatic. Nearly all surfaces of interfacing materials are indirect contact. Small areas around O-ring grooves are not in contact.However, the surfaces in those areas are so small that their separationhas an insignificant influence on the results.

The thermal results are shown in FIG. 47. Each probe location is labeledon the right and graphed on the left in a temperature vs. time plot. Thetemperature gradient plot is shown on the right at 300 seconds. Due tothe long duration inter-pulse divert coast times associated with the1000 second duty cycle it is believed that the 300 second analysis isconservative because the system cools down between pulses.

The results show that no specific area is excessively hot. NormallyO-ring temperatures and material limits are the problematic areas. Inthis application, the temperatures are low for O-rings so they do nothave any thermal problems. The material limits are based on the stressand were also determined not to be excessive. Seven probe locations arechosen for their location or temperature. The peak probe temperature isabout 450° F. at location number seven at about 225 seconds into theduty cycle, at that time a divert thrust event is occurring. After theevent occurs the interface cools down.

Example 6

A structural analysis of a number of significant components in theprototype divert system 400 are analyzed. Analyzed components includethe linkage between the divert thruster 430 and the electromechanicalactuator, the structural strength of the gas generators 418, anddeflection at the thruster interface and the O-ring interface. All ofthe components passed the structural analyses.

Example 7

A computational thermal analysis of the divert thruster 430 isperformed. The analysis determined that the divert thruster 430 is verylikely to survive the divert prototype 1000 second multi-pulse dutycycle due to the inherent 500 sec cool down time between pulses.

Illustrative Implementations

The following is a description of various implementations of thedisclosed subject matter. Each implementation may include one or more ofthe various features, characteristics, or advantages of the disclosedsubject matter. The implementations are intended to illustrate a fewaspects of the disclosed subject matter and should not be considered acomprehensive or exhaustive description of all possible implementations.

In one implementation, an attitude control system comprises: a gasgenerator including a propellant; an accumulator coupled to the gasgenerator, the accumulator being in fluid communication with the gasgenerator to allow hot gas produced by burning the propellant to flowbetween the accumulator and the gas generator; and a valve positionedbetween the gas generator and the accumulator, the valve including amain body; wherein the main body extends into the accumulator.

The valve can be an accumulator valve and the attitude control systemcan comprise a vent valve and a passage extending between the gasgenerator and the accumulator valve, wherein the vent valve movesbetween an open position where the passage is open to the outside and aclosed position where the passage is not open to the outside. Theattitude control system can comprise a valve shaft that moves between afirst position where the valve is closed and a second position where thevalve is open, the valve shaft including a ceramic matrix composite.

Pressure in the accumulator can cause hoop compression of the portion ofthe main body extending into the accumulator. The main body can includea ceramic matrix composite. The main body can include C—ZrOC or C—SiC.The attitude control system can comprise one or more thrusters coupledto the accumulator. The valve can be an accumulator valve and theattitude control system can comprise a divert valve that moves betweenan open position where the accumulator and/or the gas generator are influid communication with a divert system and a closed position where theaccumulator and/or the gas generator are not in fluid communication withthe divert system.

In another implementation, an attitude control system comprises: a gasgenerator including a propellant; an accumulator coupled to the gasgenerator, the accumulator being in fluid communication with the gasgenerator to allow hot gas produced by burning the propellant to flowbetween the accumulator and the gas generator; and a valve positionedbetween the gas generator and the accumulator, the valve including amain body made of a ceramic matrix composite.

The valve can be an accumulator valve and the attitude control systemcan comprise a vent valve and a passage extending between the gasgenerator and the accumulator valve, wherein the vent valve movesbetween an open position where the passage is open to the environmentoutside the attitude control system and a closed position where thepassage is not open to the environment outside the attitude controlsystem.

The main body can include C—ZrOC or C—SiC. The attitude control systemcan comprise a valve shaft that moves between a first position where thevalve is closed and a second position where the valve is open, the valveshaft including a ceramic matrix composite. The valve shaft can includeC—ZrOC or C—SiC. Pressure in the accumulator can cause hoop compressionof at least a portion of the main body of the valve.

The attitude control system can comprise one or more thrusters coupledto the accumulator. The valve can be an accumulator valve and theattitude control system can comprise a divert valve that moves betweenan open position where the accumulator and/or the gas generator are influid communication with a divert system and a closed position where theaccumulator and/or the gas generator are not in fluid communication withthe divert system.

In another implementation, an attitude control system comprises: a gasgenerator including a propellant; an accumulator coupled to the gasgenerator, the accumulator being in fluid communication with the gasgenerator to allow hot gas produced by burning the propellant to flowbetween the accumulator and the gas generator; and a valve positionedbetween the gas generator and the accumulator; wherein the attitudecontrol system is a low level attitude control system for a guidedmissile.

The total impulse produced by attitude control system can be no morethan 700 lbf·s. The valve can be an accumulator valve and the attitudecontrol system can comprise a vent valve and a passage extending betweenthe gas generator and the accumulator valve, wherein the vent valvemoves between an open position where the passage is open to the outsideand a closed position where the passage is not open to the outside.

The attitude control system can comprise a valve shaft that movesbetween a first position where the valve is closed and a second positionwhere the valve is open, the valve shaft including a ceramic matrixcomposite. Pressure in the accumulator can cause hoop compression of atleast a portion of the valve. The valve can comprise a main bodyincluding a ceramic matrix composite. The attitude control system cancomprise one or more thrusters coupled to the accumulator.

The valve can be an accumulator valve and the attitude control systemcan comprise a divert valve that moves between an open position wherethe accumulator and/or the gas generator are in fluid communication witha divert system and a closed position where the accumulator and/or thegas generator are not in fluid communication with the divert system.

In another implementation, a method for controlling the attitude of aflight vehicle comprises: burning propellant in a gas generator toproduce hot gas; storing the hot gas in an accumulator; and releasingthe hot gas in the accumulator through one or more thrusters to controlthe attitude of the flight vehicle.

The method can comprise extinguishing the propellant in the gasgenerator when the pressure in the accumulator reaches a set point. Theset point can be a first set point and the method can comprise ignitingthe propellant in the gas generator when a second set point is reached.The second set point can be a minimum pressure level in the accumulatoror a set amount of time that has passed since a previous event.

The method can comprise repeatedly igniting and extinguishing thepropellant in the gas generator to repeatedly pressurize the accumulatorwith the hot gas. The method can comprise burning an initial charge ofpropellant in the accumulator to pressurize the accumulator with hotgas. The method can comprise igniting the propellant in the gasgenerator for the first time with the hot gas generated by the initialcharge. The method can comprise igniting the propellant in the gasgenerator with the hot gas stored in the accumulator. The method cancomprise igniting propellant in a divert system using the hot gas in theaccumulator. The flight vehicle can be a guided missile.

In another implementation, a method for controlling the attitude of aflight vehicle comprises: burning propellant in a gas generator toproduce hot gas; storing the hot gas in an accumulator; closing a valvepositioned between the gas generator and the accumulator to prevent hotgas from flowing between the gas generator and the accumulator; andextinguishing the propellant in the gas generator.

The method can comprise releasing the hot gas in the accumulator throughone or more thrusters to control the attitude of the flight vehicle.Extinguishing the propellant in the gas generator can include opening avent valve. The method can comprise opening the valve to allow the hotgas in the accumulator to flow to the gas generator and reignite thepropellant. Opening the valve can include opening the valve when thepressure in the accumulator reaches a minimum level or a set amount oftime has passed since a previous event. Closing the valve can includeclosing the valve when the pressure in the accumulator reaches a setpoint. The flight vehicle can be a guided missile.

P1. A method for guiding a flight vehicle comprising: (a) igniting solidpropellant in a hot gas generator on the flight vehicle and generatinghot gas; (b) discharging the hot gas through at least one divertthruster on the flight vehicle; (c) extinguishing the solid propellantby rapidly depressurizing the hot gas generator; and repeating (a) and(c) at least once; wherein (b) and (c) can be performed in any order.

P2. The method of paragraph P1 comprising repeating (a) and (c) at leasttwice.

P3. The method of any one of paragraphs P1-P2 comprising igniting thesolid propellant in the hot gas generator with hot gas stored in a hotgas accumulator.

P4. The method of any one of paragraphs P1-P3 comprising igniting solidpropellant in a hot gas accumulator on the flight vehicle and ignitingthe solid propellant in the hot gas generator with hot gas from the hotgas accumulator.

P5. The method of paragraph P4 comprising igniting solid propellant inan attitude control gas generator with the hot gas from the hot gasaccumulator.

P6. The method of any one of paragraphs P1-P5 comprising igniting thesolid propellant in the hot gas generator with an igniter.

P7. The method of any one of paragraphs P1-P6 wherein rapidlydepressurizing the hot gas generator comprises opening an extinguishmentvalve.

P8. The method of any one of paragraphs P1-P7 wherein the hot gasgenerator is a first hot gas generator, and wherein (a) comprisesigniting solid propellant in a second hot gas generator on the flightvehicle and generating hot gas.

P9. The method of paragraph P8 wherein the first hot gas generator andthe second hot gas generator are part of a divert system and the firsthot gas generator and the second hot gas generator are spaced apart andpositioned opposite each other along a lengthwise axis of the divertsystem.

P10. The method of paragraph P9 wherein the first hot gas generator andthe second hot gas generator are positioned symmetrically on thelengthwise axis of the divert system.

P11. The method of any one of paragraphs P1-P10 comprising providingon-demand, multi-pulse divert thrust for at least 1000 seconds operationtime.

P12. The method of any one of paragraphs P1-P11 wherein the flightvehicle is a kill vehicle of a missile defense interceptor missile.

P13. The method of any one of paragraphs P1-P12 comprising dischargingthe hot gas through at least one attitude control thruster.

P14. A divert system for a flight vehicle comprising: a hot gasgenerator including solid propellant positioned in the hot gasgenerator; divert thrusters pneumatically linked to the hot gasgenerator; an extinguishment valve pneumatically linked to the hot gasgenerator, the extinguishment valve being movable between a closedposition where the hot gas generator is not vented and an open positionwhere the hot gas generator is vented; wherein the extinguishment valveis configured to rapidly depressurize the hot gas generator andextinguish the solid propellant when the extinguishment valve is movedfrom the closed position to the open position.

P15. The divert system of paragraph P14 wherein the hot gas generator ispneumatically separate from a propulsion rocket motor.

P16. The divert system of any one of paragraphs P14-P15 wherein the hotgas generator is a first hot gas generator, the diver system comprising:a second hot gas generator including solid propellant positioned in thesecond hot gas generator; wherein the first hot gas generator and thesecond hot gas generator are spaced apart and positioned opposite eachother along a lengthwise axis of the divert system; wherein the firsthot gas generator and the second hot gas generator are pneumaticallylinked to each other.

P17. The divert system of paragraph P16 wherein the divert thrusters arepositioned between first hot gas generator and the second hot gasgenerator.

P18. The divert system of any one of paragraphs P16-P17 wherein thefirst hot gas generator and the second hot gas generator are positionedsymmetrically on the lengthwise axis of the divert system.

P19. The divert system of any one of paragraphs P14-P18 comprising a hotgas accumulator pneumatically linked to the hot gas generator.

P20. The divert system of paragraph P19 comprising solid propellantpositioned in the hot gas accumulator.

P21. The divert system of any one of paragraphs P19-P20 comprising anattitude control hot gas generator pneumatically linked to the hot gasaccumulator.

P22. A divert system for a flight vehicle comprising: a first hot gasgenerator including solid propellant positioned in the first hot gasgenerator; a second hot gas generator including solid propellantpositioned in the second hot gas generator; divert thrusterspneumatically linked to at least one of the first hot gas generator orthe second hot gas generator; wherein the first hot gas generator andthe second hot gas generator are spaced apart and positioned oppositeeach other along a lengthwise axis of the divert system; wherein thefirst hot gas generator and the second hot gas generator arepneumatically linked to each other; wherein the first hot gas generatorand the second hot gas generator are pneumatically separate from apropulsion rocket motor; and wherein the divert thrusters are positionedbetween first hot gas generator and the second hot gas generator.

P23. The divert system of paragraph P22 comprising: an extinguishmentvalve pneumatically linked to the hot gas generator, the extinguishmentvalve being movable between a closed position where the hot gasgenerator is not vented and an open position where the hot gas generatoris vented; wherein the extinguishment valve is configured to rapidlydepressurize the hot gas generator and extinguish the solid propellantwhen the extinguishment valve is moved from the closed position to theopen position.

P24. The divert system of any one of paragraphs P22-P23 comprising a hotgas accumulator pneumatically linked to the first hot gas generator andthe second hot gas generator.

P25. The divert system of paragraph P24 comprising solid propellantpositioned in the hot gas accumulator.

P26. The divert system of any one of paragraphs P24-P25 comprising anattitude control hot gas generator pneumatically linked to the hot gasaccumulator.

P27. The divert system of any one of paragraphs P22-P26 wherein thefirst hot gas generator and the second hot gas generator are positionedsymmetrically on the lengthwise axis of the divert system.

General Terminology and Interpretative Conventions

Any methods described in the claims or specification should not beinterpreted to require the steps to be performed in a specific orderunless expressly stated otherwise. Also, the methods should beinterpreted to provide support to perform the recited steps in any orderunless expressly stated otherwise.

Certain features described in the context of separate implementationscan also be implemented in combination in a single implementation.Conversely, various features that are described in the context of asingle implementation can also be implemented in multipleimplementations separately or in any suitable subcombination. Moreover,although features may be described above in certain combinations andeven initially claimed as such, one or more features from a claimedcombination can be excised from the combination, and the claimedcombination may be directed to a subcombination or variation of asubcombination.

Articles such as “the,” “a,” and “an” can connote the singular orplural. Also, the word “or” when used without a preceding “either” (orother similar language indicating that “or” is unequivocally meant to beexclusive—e.g., only one of x or y, etc.) shall be interpreted to beinclusive (e.g., “x or y” means one or both x or y).

The term “and/or” shall also be interpreted to be inclusive (e.g., “xand/or y” means one or both x or y). In situations where “and/or” or“or” are used as a conjunction for a group of three or more items, thegroup should be interpreted to include one item alone, all the itemstogether, or any combination or number of the items.

The terms have, having, include, and including should be interpreted tobe synonymous with the terms comprise and comprising. The use of theseterms should also be understood as disclosing and providing support fornarrower alternative implementations where these terms are replaced by“consisting” or “consisting essentially of.”

Unless otherwise indicated, all numbers or expressions, such as thoseexpressing dimensions, physical characteristics, and the like, used inthe specification (other than the claims) are understood to be modifiedin all instances by the term “approximately.” At the very least, and notas an attempt to limit the application of the doctrine of equivalents tothe claims, each numerical parameter recited in the specification orclaims which is modified by the term “approximately” should be construedin light of the number of recited significant digits and by applyingordinary rounding techniques.

All disclosed ranges are to be understood to encompass and providesupport for claims that recite any subranges or any and all individualvalues subsumed by each range. For example, a stated range of 1 to 10should be considered to include and provide support for claims thatrecite any and all subranges or individual values that are betweenand/or inclusive of the minimum value of 1 and the maximum value of 10;that is, all subranges beginning with a minimum value of 1 or more andending with a maximum value of 10 or less (e.g., 5.5 to 10, 2.34 to3.56, and so forth) or any values from 1 to 10 (e.g., 3, 5.8, 9.9994,and so forth), which values can be expressed alone or as a minimum value(e.g., at least 5.8) or a maximum value (e.g., no more than 9.9994).

All disclosed numerical values are to be understood as being variablefrom 0-100% in either direction and thus provide support for claims thatrecite such values (either alone or as a minimum or a maximum—e.g., atleast <value> or no more than <value>) or any ranges or subranges thatcan be formed by such values. For example, a stated numerical value of 8should be understood to vary from 0 to 16 (100% in either direction) andprovide support for claims that recite the range itself (e.g., 0 to 16),any subrange within the range (e.g., 2 to 12.5) or any individual valuewithin that range expressed individually (e.g., 15.2), as a minimumvalue (e.g., at least 4.3), or as a maximum value (e.g., no more than12.4).

The terms recited in the claims should be given their ordinary andcustomary meaning as determined by reference to relevant entries inwidely used general dictionaries and/or relevant technical dictionaries,commonly understood meanings by those in the art, etc., with theunderstanding that the broadest meaning imparted by any one orcombination of these sources should be given to the claim terms (e.g.,two or more relevant dictionary entries should be combined to providethe broadest meaning of the combination of entries, etc.) subject onlyto the following exceptions: (a) if a term is used in a manner that ismore expansive than its ordinary and customary meaning, the term shouldbe given its ordinary and customary meaning plus the additionalexpansive meaning, or (b) if a term has been explicitly defined to havea different meaning by reciting the term followed by the phrase “as usedin this document shall mean” or similar language (e.g., “this termmeans,” “this term is defined as,” “for the purposes of this disclosurethis term shall mean,” etc.). References to specific examples, use of“i.e.,” use of the word “invention,” etc., are not meant to invokeexception (b) or otherwise restrict the scope of the recited claimterms. Other than situations where exception (b) applies, nothingcontained in this document should be considered a disclaimer ordisavowal of claim scope.

The subject matter recited in the claims is not coextensive with andshould not be interpreted to be coextensive with any implementation,feature, or combination of features described or illustrated in thisdocument. This is true even if only a single implementation of thefeature or combination of features is illustrated and described.

Joining or Fastening Terminology and Interpretative Conventions

The term “coupled” means the joining of two members directly orindirectly to one another. Such joining may be stationary in nature ormovable in nature. Such joining may be achieved with the two members orthe two members and any additional intermediate members being integrallyformed as a single unitary body with one another or with the two membersor the two members and any additional intermediate member being attachedto one another. Such joining may be permanent in nature or alternativelymay be removable or releasable in nature.

The term “coupled” includes joining that is permanent in nature orreleasable and/or removable in nature. Permanent joining refers tojoining the components together in a manner that is not capable of beingreversed or returned to the original condition. Releasable joiningrefers to joining the components together in a manner that is capable ofbeing reversed or returned to the original condition.

Drawing Related Terminology and Interpretative Conventions

The drawings are intended to illustrate implementations that are bothdrawn to scale and/or not drawn to scale. This means the drawings can beinterpreted, for example, as showing: (a) everything drawn to scale, (b)nothing drawn to scale, or (c) one or more features drawn to scale andone or more features not drawn to scale. Accordingly, the drawings canserve to provide support to recite the sizes, proportions, and/or otherdimensions of any of the illustrated features either alone or relativeto each other. Furthermore, all such sizes, proportions, and/or otherdimensions are to be understood as being variable from 0-100% in eitherdirection and thus provide support for claims that recite such values orany and all ranges or subranges that can be formed by such values.

Spatial or directional terms, such as “left,” “right,” “front,” “back,”and the like, relate to the subject matter as it is shown in thedrawings and/or how it is commonly oriented during manufacture, use, orthe like. However, it is to be understood that the described subjectmatter may assume various alternative orientations and, accordingly,such terms are not to be considered as limiting.

INCORPORATION BY REFERENCE

The entire contents of each of the documents listed below areincorporated by reference into this document (the documents below arecollectively referred to as the “incorporated documents”). If the sameterm is used in both this document and one or more of the incorporateddocuments, then it should be interpreted to have the broadest meaningimparted by any one or combination of these sources unless the term hasbeen explicitly defined to have a different meaning in this document. Ifthere is an inconsistency between any of the following documents andthis document, then this document shall govern. The incorporated subjectmatter should not be used to limit or narrow the scope of the explicitlyrecited or depicted subject matter.

-   -   U.S. Prov. App. No. 62/670,685, titled “Flight Vehicle Divert        System,” filed on 11 May 2018.    -   U.S. Pat. No. 9,927,217 (application. Ser. No. 14/847,820),        titled “Attitude Control System,” filed on 8 Sep. 2015, issued        on 27 Mar. 2018.    -   U.S. patent application Ser. No. 15/488,267, titled “Hot Gas        Thruster,” filed on 14 Apr. 2017.    -   U.S. patent application Ser. No. 14/875,424, titled “Attitude        Control System,” filed on 5 Oct. 2015.

The invention claimed is:
 1. A method for controlling the trajectory ofa flight vehicle comprising: (a) igniting solid propellant in a firsthot gas generator disposed in a first case and a second hot gasgenerator disposed in a second case on the flight vehicle and generatinghot gas, the solid propellant in the first hot gas generator and thesecond hot gas generator burning at the same time, wherein the secondcase is spaced apart from the first case along a lengthwise axis; (b)discharging the hot gas through at least one divert thruster on theflight vehicle; (c) extinguishing the solid propellant by rapidlydepressurizing the first hot gas generator and the second hot gasgenerator; igniting solid propellant in a hot gas accumulator on theflight vehicle and igniting the solid propellant in the first hot gasgenerator and/or the second hot gas generator with hot gas from the hotgas accumulator, wherein the hot gas accumulator is configured to storehot gas; and repeating (a) and (c) at least once; wherein (b) and (c)can be performed in any order.
 2. The method of claim 1 comprisingrepeating (a) and (c) at least twice.
 3. The method of claim 1comprising igniting the solid propellant in the first hot gas generatorand/or the second hot gas generator with hot gas stored in a hot gasaccumulator.
 4. The method of claim 1 comprising igniting solidpropellant in an attitude control gas generator with the hot gas fromthe hot gas accumulator.
 5. The method of claim 1 comprising ignitingthe solid propellant in the first hot gas generator and/or the secondhot gas generator with an igniter.
 6. The method of claim 1 whereinrapidly depressurizing the first hot gas generator and the second hotgas generator comprises opening an extinguishment valve.
 7. The methodof claim 1 wherein the first hot gas generator and the second hot gasgenerator are part of a divert system and the first hot gas generatorand the second hot gas generator are spaced apart and positionedopposite each other along a lengthwise axis of the divert system.
 8. Themethod of claim 1 comprising providing on-demand, multi-pulse divertthrust for at least 1000 seconds operation time.
 9. The method of claim1 wherein the flight vehicle is a kill vehicle of a missile defenseinterceptor missile.
 10. A divert system for a flight vehiclecomprising: a first hot gas generator including solid propellantpositioned in the first hot gas generator; a second hot gas generatorincluding solid propellant positioned in the second hot gas generator;divert thrusters on the flight vehicle pneumatically linked to at leastone of the first hot gas generator or the second hot gas generator; andan extinguishment valve pneumatically linked to the first hot gasgenerator and the second hot gas generator, the extinguishment valvebeing movable between a closed position where the hot gas generator isnot vented and an open position where the hot gas generator is vented;wherein the extinguishment valve is configured to rapidly depressurizethe hot gas generator and extinguish the solid propellant when theextinguishment valve is moved from the closed position to the openposition; wherein the first hot gas generator and the second hot gasgenerator are spaced apart along a lengthwise axis of the divert systemand positioned opposite each other along the lengthwise axis; whereinthe first hot gas generator and the second hot gas generator arepneumatically linked to each other; a hot gas accumulator pneumaticallylinked to the first hot gas generator and/or the second hot gasgenerator; and solid propellant positioned in the hot gas accumulator.11. The divert system of claim 10 wherein the first hot gas generatorand the second hot gas generator are pneumatically separate from anypropulsion rocket motors on the flight vehicle.
 12. The divert system ofclaim 10 wherein the divert thrusters are positioned between first hotgas generator and the second hot gas generator.
 13. The divert system ofclaim 10 comprising an attitude control hot gas generator pneumaticallylinked to the hot gas accumulator.
 14. A divert system for a flightvehicle comprising: a first hot gas generator including solid propellantpositioned in the first hot gas generator; a second hot gas generatorincluding solid propellant positioned in the second hot gas generator;divert thrusters pneumatically linked to at least one of the first hotgas generator or the second hot gas generator; wherein the first hot gasgenerator and the second hot gas generator are spaced apart along alengthwise axis of the divert system and positioned opposite each otheralong the lengthwise axis; wherein the first hot gas generator and thesecond hot gas generator are pneumatically linked to each other; whereinthe first hot gas generator and the second hot gas generator arepneumatically separate from any propulsion rocket motors on the flightvehicle; wherein the divert thrusters are positioned between the firsthot gas generator and the second hot gas generator along the lengthwiseaxis of the divert system; a hot gas accumulator pneumatically linked tothe first hot gas generator and/or the second hot gas generator; andsolid propellant positioned in the hot gas accumulator.
 15. The divertsystem of claim 14 comprising: an extinguishment valve pneumaticallylinked to the first hot gas generator and/or the second hot gasgenerator, the extinguishment valve being movable between a closedposition where the first hot gas generator and/or the second hot gasgenerator is not vented and an open position where the first hot gasgenerator and/or the second hot gas generator is vented; wherein theextinguishment valve is configured to rapidly depressurize the first hotgas generator and/or the second hot gas generator and extinguish thesolid propellant when the extinguishment valve is moved from the closedposition to the open position.
 16. The divert system of claim 14comprising an attitude control hot gas generator pneumatically linked tothe hot gas accumulator.